Abstract: Crack propagation is an important concern in the design of aircraft composite fuselage and wing panels. However, numerical simulation of crack propagation is computationally expensive. This work proposes combining high-fidelity analysis model with low-fidelity model to calculate the crack propagation constraint in the design optimization process. Correction response surfaces are employed to relate the high-fidelity models to the low-fidelity models. Four different forms of correction response surface methods are explored and their prediction capabilities are compared. The multi-fidelity approach is found to be more accurate than single-fidelity response surface method at the same computational cost.

Abstract: 1. Introduction. 2. Preliminary Estimate of Take-Off Weight. 3. Wing Loading Selection. 4. Main Wing Design. 5. Fuselage Design. 6. Horizontal and Vertical Tail Design. 7. Engine Selection. 8. Take-Off and Landing. 9. Enhanced Lift Design. 10. Structural Design and Material Selection. 11. Static Stability and Control. 12. Cost Estimate. 13. Design Summary and Trade Study.

Abstract: A vehicle including a fuselage, at least one lift-producing propeller carried by the fuselage on each side of its transverse axis a pilot's compartment formed in the fuselage between the lift-producing propellers and substantially aligned with the longitudinal axis, and a pair of payload bays formed in the fuselage between the lift-producing propellers and on opposite sides of the pilot's compartment. Many variations are described enabling the vehicle to be used not only as a VTOL vehicle, but also as a multi-function utility vehicle for performing many diverse functions including hovercraft and ATV functions.

Abstract: An integrated and modular high-speed aircraft (200) and method of design and manufacture. The aircraft (200) can have a supersonic or near-sonic cruise Mach number. In one embodiment, the aircraft (200) can include an aft body integrated with a delta wing (204) and a rearwardly-tapering fuselage (202) to define a smooth forward-to-rear area distribution. A propulsion system (206), including an engine (216), inlet (220), and exhaust nozzle (222) can be integrated into the aft body to be at least partially hidden behind the wing (204). In one embodiment, the entrance of the inlet can be positioned beneath the wing (204), and the exit of the nozzle (222) can be positioned at or above the wing (204). An S-shaped inlet duct (221) can deliver air to the aft-mounted integrated engine.

Abstract: An aerial vehicle (10) including a toroidal fuselage (12) having a longitudinal axis (14), and a duct (16) extending along the longitudinal axis (14) between a leading edge (18) and a trailing edge (20) of the fuselage (12), first and second counter-rotating, variable pitch rotor assemblies (22, 24) coaxially mounted within the duct (16) of the fuselage (12), and at least one canard wing (26) secured to the toroidal fuselage (12) and having a leading edge (28) positioned out of the duct (16) of the fuselage (12) and axially forward of the leading edge of the fuselage (12), wherein at least a portion of the canard wing (26) comprises a control surface having a variable angle of attack. The invention provides an aerial vehicle (10) that can take-off and land vertically, hover for extended periods of time over a fixed spatial point, and operated in confined areas. The aerial vehicle (10) also has the ability to transition between a hover and high speed forward flight.

Abstract: A miniature, unmanned aircraft for acquiring and/or transmitting data, capable of automatically maintaining desired airframe stability while operating by remote directional commands. The aircraft comprises a fuselage and a wing, a piston engine and propeller, a fuel supply, at least one data sensor and/or radio transceiver, a microprocessor disposed to manage flight, a radio transceiver for receiving remotely generated flight direction commands, a GPS receiver, a plurality of control surfaces and associated servomechanisms, for controlling flight stabilization and direction, roll, pitch, yaw, velocity, and altitude sensors. The microprocessor uses roll, pitch, yaw, and altitude data to control attitude and altitude of the aircraft automatically, but controls flight direction solely based on external commands. The aircraft does not exceed fifty-five pounds.

Abstract: Initial results of an investigation towards finding an efficient non-cylindrical fuselage configuration for a conceptual blended-wing-body flight vehicle were presented. A simplified 2-D beam column analysis and optimization was performed first. Then a set of detailed finite element models of deep sandwich panel and ribbed shell construction concepts were analyzed and optimized. Generally these concepts with flat surfaces were found to be structurally inefficient to withstand internal pressure and resultant compressive loads simultaneously. Alternatively, a set of multi-bubble fuselage configuration concepts were developed for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls. An outer-ribbed shell was designed to prevent buckling due to external resultant compressive loads. Initial results from finite element analysis appear to be promising. These concepts should be developed further to exploit their inherent structurally efficiency.

Abstract: The present invention provides an unmanned airborne reconnaissance vehicle having a fuselage, a forward wing pair and a rearward wing pair vertically separated by a gap and staggered fore and aft therebetween such that a general biplane configuration is formed. The present invention provides a pair of wing tip plates for joining the wing tips of the forward and rearward wings. The unmanned airborne reconnaissance vehicle of the present invention includes a power plant to propel the vehicle through the air and a generally T-shaped tail having a vertical stabilizer including a rudder and a full span elevator.

Abstract: A 25-ft/s vertical drop test of a composite fuselage section was conducted with two energy-absorbing seats occupied by anthropomorphic dummies to evaluate the crashworthy features of the fuselage section and to determine its interaction with the seats and dummies. The 5-ft. diameter fuselage section consists of a stiff structural floor and an energy-absorbing subfloor constructed of Rohacel foam blocks. The experimental data from this test were analyzed and correlated with predictions from a crash simulation developed using the nonlinear, explicit transient dynamic computer code, MSC.Dytran. The anthropomorphic dummies were simulated using the Articulated Total Body (ATB) code, which is integrated into MSC.Dytran.

Abstract: The results are presented of a study concerned with the prediction of the airflow noise transmitted through an element of the fuselage structure: a double panel of finite extent that consists of a pair of thin elastic plates containing a light insulating material separated from the inner skin by an air gap. This configuration is representative of typical compound sidewalls in large commercial aircraft. A solution based on modal coupling is obtained and validated by comparisons with other solutions on various test cases. A physical interpretation is given for the calculated vibroacoustic response of a double partition system excited by a turbulent boundary layer, and the effect of an air gap between the insulation facing bag and the trim panel is analyzed. It is shown that the levels of the inwardly radiated sound power are mainly determined by the contribution of the first skin panel-controlled mode, and the added damping effect due to the insulating material has little effect below this resonance. To achieve sound reduction in the very low-frequency domain, the performance of various active control strategies are examined and compared. It is found that the most efficient strategy is the suppression of the low-order skin panel structural modes. However, we note that significant reductions in the sound power radiated can also he achieved by the active suppression of the low-order structural modes of the trim panel.

Abstract: : A 30-ft/s vertical drop test of a 10-ft-long fuselage section of a Boeing 737 (B737) aircraft was conducted in October of 1999 at the Federal Aviation Administration (FAA) William J Hughes Technical Center, Atlantic City International Airport, NJ This test was performed to evaluate the structural integrity of a conformable auxiliary fuel tank mounted beneath the fuselage floor and to determine its effect on the structural response of the airframe A second drop test of a similar B737 fuselage section was conducted in November of 2000 in which two different overhead stowage bins were evaluated These tests present an invaluable opportunity to evaluate the capabilities of computational tools for crash simulation through analytical and experimental correlation To perform this evaluation, a full-scale three-dimensional finite element model of the fuselage section was developed A crash simulation was conducted using the explicit, nonlinear transient dynamic code, MSCDytranTM For the initial simulation, structural deformation and floor-level acceleration responses were generated and correlated with experimental data obtained during the drop test of the B737 fuselage section with the auxiliary fuel tank The focus of the follow-on simulation was to develop pretest predictions of the fuselage and overhead bin responses for correlation with data from the vertical drop test of the second B737 fuselage section An assessment of the accuracy of the pretest predictions was made and model improvements were suggested Several of the model improvements were implemented and the effects of the changes on model accuracy were evaluated

Abstract: An air vehicle, such as a manned or unmanned air vehicle, has a fuselage, a rotor/scissors wing, and a scissors wing. At helicopter mode, the rotor/scissors wing rotates to make the air vehicle fly like a helicopter to achieve vertical and/or short take-off and landing, hovering, and low speed flying. At airplane mode, the rotor/scissors wing and scissors wing form a scissors wings configuration to maximize the air vehicle's flying efficiency at a wide range of speed and flying conditions by adjusting the yaw angle of the rotor/scissors wing and scissors wing. During the conversion from helicopter mode to airplane mode, the scissors wing generates lift to offload the rotating rotor/scissors wing and eventually the offloaded rotor/scissors wing's rotating speed is slowed and stopped so that the rotor/scissors wing can be locked at a specific position and the conversion can be achieved. In a reverse order, the air vehicle can convert from airplane mode to helicopter mode. Either turbofan or turbojet engine, or turboshaft/turbofan convertible engine can be used to power the air vehicle.

Abstract: A vertical takeoff and landing aircraft ( 10 ) is provided including an aircraft fuselage ( 12 ). A plurality of hubs ( 20 ) are mechanically coupled to the fuselage ( 12 ) and are rotated by at least one engine ( 22 ). A plurality of tandem rotor/wings ( 14 ) are mechanically coupled to the plurality of hubs ( 20 ) and propel and lift the aircraft fuselage ( 12 ). A transitional lift wing ( 16 ) is mechanically coupled to the fuselage ( 12 ) and enables lift on the fuselage ( 12 ) during off-loading lift of the plurality of tandem rotor/wings ( 14 ). A main controller ( 24 ) is coupled to the plurality of tandem rotor/wings ( 14 ) and switches the plurality of tandem rotor/wings ( 14 ) between a vertical lift mode and a fixed wing mode. A method of performing the same is also provided.

Abstract: Recent advances in computational speed have made aircraft and spacecraft crash simulations using an explicit, nonlinear, transient-dynamic,e niteelement analysiscodemore feasible.This paperdescribes thedevelopment of a simplelanding-gearmodel, which accurately simulatestheenergy absorbed by thegearwithoutaddingsubstantial 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: A 10-foot long fuselage section from a Boeing 737-100 was dropped from a height of 14 feet, generating a final velocity at impact of 30 feet per second. The fuselage section was configured to simulate the load density at the maximum takeoff weight condition. The final weight of 8870 pounds included cabin seats, dummy occupants, overhead stowage bins with contents, and cargo compartment luggage. The fuselage section was instrumented with strain gages, accelerometers, and high-speed cameras. The fuselage section sustained severe deformation of the cargo compartment. The cargo compartment luggage influenced the manner in which the fuselage crushed affecting the g forces experienced by the fuselage section and the pulse duration. The seat tracks experiences a vertical impact pulse of 15 g and a pulse duration of approximately 135 milliseconds.

Abstract: Device for attaching an engine to an aircraft. The attachment device (10) is inserted between the engine (14) and a pylon (12) designed to be attached to an aircraft structure, such as a wing or fuselage section. The attachment device (10) generally includes two brackets (18a, 18b) joined together and secured to the pylon (12) and two pairs of rods (20a, 20b, 22a, 22b) inserted between the two brackets (18a, 18b) and a part (24) of the structure of the engine (14). Thus the integrity of the link between the pylon (16) and the engine (14) is preserved in the event of failure of any of the parts which constitute the device, without there being any need for a backup attachment system.

Abstract: Flow distortion is a major issue in the measurement of wind turbulence with gust probes mounted on a nose boom, at the radome, or under the wing of research aircraft. In this paper, the effects both of the propellers of a turboprop aircraft and of the aircraft vortex system on the pressure and wind velocity measurements near the nose of the aircraft are examined. It is shown that, for a turboprop aircraft, the sensors mounted near the nose are affected directly (slipstream) or indirectly (lift increase) by the propellers. The propeller effects are more significant for pressure sensors located ahead of the propellers on the fuselage and are less significant for the small local flow angles measured at the nose of the aircraft. The first case is clearly realized during in-flight calibration maneuvers performed by a turboprop aircraft. A major flow distortion, which seriously affects the vertical wind velocity measurements near the nose of an aircraft, is the upwash induced mainly by the wingbound vortex. Also, low energy of the vertical wind component in the inertial subrange for scales larger than the fuselage diameter is usually observed in aircraft measurements. This is shown to be the result of not taking into account the decrease of the upwash correction with eddy frequency (or no need for such a correction in the inertial subrange) caused by the aerodynamic delay and the response of the wing vortex to turbulence. The level of energy in the inertial subrange of the vertical wind component is significant because it is commonly used for the estimation of the dissipation rate of turbulence kinetic energy. A method to estimate this frequency variable correction and correct the spectra or the time series of the estimated vertical wind component is described. Data from low-level flight legs with a Twin Otter aircraft show that this correction may result in about a 20% correction of the variance of the vertical wind component and a 5% correction of the vertical turbulent fluxes.

Abstract: An vertical take-off and landing (VTOL) and horizontal flight (HF) aircraft has a fuselage with wings providing lift during horizontal flight, a rotor on the nose of the fuselage that is stowable during horizontal flight, a first pair of ducted fan propellers on the wings, a second pair of tiltable ducted fan propellers adjacent to the tail, a hydraulic system powered by the main engine and generating propulsive hydraulic power for the rotor system and the first and second pairs of propellers. A control system for the hydraulic system provides propulsive hydraulic power to the rotor system and to the second pair of propellers during vertical take-off and landing and not to the first pair of propellers during vertical take-off and landing and provides propulsive power to the first and second pairs of propellers during horizontal flight and not to the rotor system during horizontal flight.

Abstract: A miniature unmanned aircraft which uses remotely controlled model aircraft components and technology, and has on-board automatic “on-the-fly” fuel and air mixture adjustment enabling high altitude flight. The aircraft, which may have conventional fuselage, wing, reciprocating piston engine and radio frequency operated controls, also has sensors for sensing atmospheric pressure, atmospheric temperature, engine crankshaft rotational speed, engine temperature, and exhaust temperature. A microprocessor aboard the aircraft receives inputs from the sensors and controls at least one servo to adjust fuel and air mixture according to preprogrammed look-up tables and equations to operate the engine at appropriate fuel-to-air ratios for the altitude and other operating conditions.

Abstract: A stairway module for use within a fuselage of an aircraft for enabling access to an overhead storage or sleeping area within the aircraft. One preferred embodiment discloses a spiral stairway module having a footprint no larger than a standard lavatory module. Another embodiment includes a stairway module having a pivotally mounted stair section which can be moved between operable and inoperable positions. Both embodiments provide a mid-level platform changing area. All of the embodiments are extremely compact and allow either fore-to-aft access or athwartship access to the overhead sleeping or storage area in the fuselage of the aircraft.

Abstract: An air vehicle, such as an aircraft, an unmanned air vehicle, a missile, or an aero bomb that has a fuselage and two main wings each of which has a left side wing and a right side wing Both of the main wings are rotatably mounted on the fuselage via one or two pivots or hollow turrets so that both of them can be yawed during flight to optimize flying efficiency under various flying conditions

Abstract: The TANGO Composite Fuselage is one of four large technology platforms that constitute the TANGO program (Technology Application to the Near-Term Business Goals and Objectives of the Aerospace Industry). The TANGO program itself is a large consortium of 34 airframe manufacturers, equipment and component suppliers, academic institutes and universities. The primary objectives of TANGO are to develop and mature technologies capable of delivering 20% cost and weight savings. over current state of the art aircraft structures. The TANGO Composite Fuselage is the first opportunity, in Europe, to develop a large scale test platform representative of a single aisle commercial airliner fuselage. The platform uses mature and emerging composite manufacturing technologies and assembles them in a full scale fuselage barrel test specimen. The Composite Fuselage platform involves 9 partners from 7 European countries (Belgium, France, Germany, Greece, Holland, Italy, Spain) and is partially funded by the European Union under the Key Action Aeronautics of the GROWTH RTD. The partners include airframe manufacturers, research institutes and academia who are all involved in the full integration of the platform and the development of the final test specimen. The platform consists of a number of specific work packages which cover the 4 years of the TANGO programme. The first phase, of about a year, includes the definition of the fuselage, the test programme and the philosophies to be used for the design, manufacturing and assembly of a composite fuselage. The second phase concentrates on the design of the fuselage section, the manufacturing and the final assembly. This phase is approximately 2 years. The third and final phase is the full scale test of the fuselage section. The test programme is expected to take a year. The technologies in the Composite Fuselage platform include Automatic Fiber Placement (AFP), Resin Film Infusion (RFI), Liquid Resin Infusion (LRI), Resin Transfer Molding (RTM) and thermoplastic technologies. In addition, traditional prepreg tape laying will be applied. The materials to be used will include braided fibres, woven fabrics, non-crimp fabrics and thermoset and thermoplastic prepreg tapes. The results of the final test will be used in the evaluation of the Composite Fuselage with respect to the stated TANGO goals of 20% weight and cost savings.

Abstract: An aircraft includes a fuselage and a fixed flying surface having two wings, and a cradle-shaped ventral fairing at the crossover of the latter with the flying surface and provided with two longitudinal flanges climbing laterally along the fuselage and being provided with an opening for the passage with play of the corresponding wing to form a first peripheral slot around the ventral fairing and a second peripheral slot around each wing. An elongate seal is fixed to the periphery of the ventral fairing and to the periphery of each opening, so at to shut off the first and second slots, respectively. The seal includes a longitudinal elastic end lip applied, via the inner face of its free end, respectively against the fuselage or the flying surface and a longitudinal leaktightness member, carried by the inner face of the seal set back with respect to the end lip and also able to be applied against the fuselage or flying surface, respectively. The additional longitudinal leaktightness member is radially elastic bead.

Abstract: Optimal tiltrotor flight paths in the event of a single engine failure during short takeoff operations are investigated. A vertical plane rigid-body model is used that has as state variables aircraft position, body components of aircraft velocity, pitch angle and rate, and rotor angular speed. The control variables are the thrust coefficient of one rotor, the pilot's longitudinal stick displacement, and the nacelle angle. The model uses both parameters and aerodynamic data of the XV-15 research aircraft. The tabular aerodynamic data are interpolated with smooth functions. Tiltrotor flights after engine failure are formulated as nonlinear optimal control problems. Both continued takeoff and rejected takeoff flight after an engine failure are studied. Performance indices are selected to minimize runway length, subject to various constraints from safety considerations and tiltrotor performance limitations. These optimal control problems are parameterized via collocation into parameter optimizations for numerical solutions. Extensive numerical solutions are obtained, and sensitivity analyses are conducted to examine effects of model uncertainties

Abstract: The paper presents a time-domain method for computation of sound radiation from aircraft engine sources to the far-field The effects of nonuniform flow around the aircraft and scattering of sound by fuselage and wings are accounted for in the formulation Our approach is based on the discretization of the inviscid flow equations through a collocation form of the Discontinuous Galerkin spectral element method An isoparametric representation of the underlying geometry is used in order to take full advantage of the spectral accuracy of the method Largescale computations are made possible by a parallel implementation based on message passing Results obtained for radiation from an axisymmetric nacelle alone are compared with those obtained when the same nacelle is installed in a generic conguration, with and without a wing

Abstract: The invention concerns a remote-controlled flying machine, in particular for surveillance and inspection, capable of hovering and comprising a spherical open-worked resistant shroud integral with a cylindrical fairing wherein rotates a propeller powered by an engine housed in a fuselage secured to the fairing with radial arms and straightening vanes.