Abstract: Buffeting is an aeroelastic phenomenon that plagues high performance aircraft, especially those with twin vertical tails. Unsteady vortices emanate from wing/fuselage leading edge extensions when these aircraft maneuver at high angles of attack. These aircraft are designed such that the vortices shed while maneuvering at high angles of attack and improve the lift-to-drag ratio of the aircraft. With proper placement and sizing of the vertical tails, this improvement may be maintained without adverse effects to the tails. However, there are tail locations and angles of attack where these vortices burst and immerse the vertical tails in their wake inducing severe structural vibrations. The resulting buffet loads and severe vertical tail response became an airframe life and maintenance concern as life cycle costs increased. Several passive methods have been investigated to reduce the buffeting of these vertical tails with limited success. As demonstrated through analyses, wind-tunnel investigations, and full-scale ground tests, active control systems offer a promising solution to alleviate buffet induced strain and increase the fatigue life of vertical tails. A collaborative research project including the United States, Canada, and Australia is in place to demonstrate active buffet load alleviation systems on military aircraft. The present paper provides details on this collaborative project and other research efforts to reduce the buffeting response of vertical tails in fighter aircraft.

Abstract: A fixed-wing four-to-six-seat light aircraft that can be easily converted to a roadway vehicle within minutes by a single person in the field, comprising a one-piece wing center panel with foldable wing tips on each sides. The whole wing unit is then rotatably mounted on top of the fuselage. The aircraft features a conventional front-engine-and-propeller lay-out, with a foldable tail section for convenient roadability and garageability. All the wheels are retractable in flight, and it has a long-span, high aspect-ratio wing for exceptional climb and cruise efficiency. The vehicle has a low ride-height with a low center of gravity, four wheels with independent suspension, nose-height leveling for take-off and landing, and anti-sway mechanism for good ground handling ability. Ground propulsion is by automotive-style transmission driving the rear wheels or by hydraulic motors in the all-wheel retractable version.

Abstract: Textile reinforced composite structural elements that have been developed in the NASA ACT Program are discussed. Included are braided fuselage frames and window-belt reinforcements, woven/stitched lower fuselage side panels, stitched multiaxial warp knit wing skins, and braided wing stiffeners. In addition, low-cost processing concepts such as resin transfer molding (RTM), resin film infusion (RFI), and vacuum-assisted resin transfer molding (VARTM) are discussed. Process modeling concepts to predict resin flow and cure in testile preforms are also discussed.

Abstract: Aircraft for carrying passengers and/or freight based on a known aircraft design with a fuselage having a nose section, center section, and tail section and with airfoils mounted on the center section near the center of gravity of the aircraft and calculated for the required lift and with vertical and/or horizontal stabilizers located in the vicinity of the tail section for creating stabilizing and steering moments, with the fuselage of said aircraft being stretched by means of an additional section inserted between the nose section and the center of gravity of the aircraft model in order to increase the carrying capacity of the aircraft based on predetermined known aircraft designs, and with the additional section, as viewed in the direction of flight, being equipped forward of the airfoils of the aircraft model that serve as the main airfoils with airfoils that serve as additional airfoils, with previously designed and calculated airfoils of suitable size from a known aircraft design being used as the additional airfoils.

Abstract: A fuselage is usually associated with the use of a delta wing on an actual aircraft. It is also in many cases necessary to use a centerbody of some shape in tunnel tests of pure delta wings. The present paper describes the e ow physics causing the experimentally observed large effect of a fuselage on delta-wing vortex breakdown.

Abstract: An aerial refueling system includes a pod assembly which is removably mounted for translation on an elongated pylon secured to the underside of an aircraft fuselage, whereby the pod assembly is moved relatively forward on the fuselage for stowage and relatively aftward on the fuselage for refueling operations. The pod assembly includes a pivotable, telescoping boom having control surfaces by which to aerodynamically position the boom's free end relative to the pod assembly. A coupling preferably establishes fluid communication between the boom and a fuel tank within the aircraft fuselage only when the pod assembly assumes the relatively-aftward refueling position on the pylon. A ram-air turbine on the forward end of the pod assembly provides all necessary power for the pod assembly, with wireless remote operation of all pod assembly functions preferably used to minimize the extent of alterations when installing the aerial refueling system on the aircraft.

Abstract: An aircraft has a fuselage substantially designed as an aerostatic lifting body and combined lifting and propelling devices joined to the fuselage, provided with propellers and forming driving units which can tilt between a lifting position, in which the plane of rotation of the respective propeller is substantially horizontal and the driven shaft of the associated drive that drives the propeller shaft is substantially vertical, and a propelling position in which the plane of rotation of the respective propeller is substantially vertical and the driven shaft of the associated drive that drives the propeller shaft is substantially horizontal. The plane of rotation of the propeller can swivel around the driven shaft of the associated drive that drives the propeller shaft.

Abstract: A jet aircraft has a generally conical front fuselage section, a cylindrical intermediate fuselage section defining a passenger compartment, a generally conical aft fuselage section, and a single vertical stabilizer. The aircraft's propulsion engines are mounted on pylons on the conical aft fuselage section with the air inlets thereof disposed entirely within a rearward projection of the lateral cross section of the intermediate fuselage section thereby to preclude the ingestion of foreign objects into the engines while minimizing the effect of boundary layer airflow. The exhaust nozzles extend rearwardly past the vertical stabilizer to minimize side line noise.

Abstract: A mechanism for quickly removing and installing the free wings and/or the tail booms of a freewing aircraft to the fuselage. The fuselage includes a free wing cross tube extending transversely through the fuselage at the spanwise axis, and each of the left and right free wings includes a support tube disposed therein, also along the spanwise axis, with a portion of the wing structural tube being received within the fuselage cross tube. A quick-release pin is insertable in holes disposed in both the structural tube and the cross tube when the holes are placed in alignment. A second cross tube extends transversely through the rear end of the fuselage at the tail boom pivot axis. At least a portion of each tail boom member is disposed in surrounding relationship to the cross tube. In a preferred embodiment, a quick-release pin is insertable in holes disposed in both the tail boom member portion and the cross tube when the holes are placed in alignment.

Abstract: An aircraft structure has an arrangement of aircraft components that provide inherent directional stability for a flight vehicle throughout an angle-of-attack range, even at very high angles-of-attack where conventional means of stabilization are ineffective. Components attached to an aircraft fuselage include a wing, horizontal stabilizers and vertical stabilizers. The wing is mounted forward of the horizontal stabilizers and is carried high on the fuselage. The horizontal stabilizer is mounted toward the rear of the aircraft and is attached near the bottom of the fuselage. The wing and horizontal stabilizers are joined on either side of the aircraft by forwardly sweeping aerodynamically shaped surfaces serving as the vertical stabilizers. The inclination of the vertical stabilizers preferably ranges from 45 degrees (top edge canted outboard) to 90 degrees (panels vertical). Preferably, the surface area of the vertical stabilizers is concentrated aft such that the aerodynamic center of the vertical stabilizers is located behind the center-of-gravity of the aircraft.

Abstract: A closed, elastic cylindrical shell is used as a simple model of the aircraft’s fuselage. For reducing the vibration of the fuselage and the subsequent interior sound pressure due to the propellers, several dynamic absorbers are attached to the shell. In the paper, we analytically derive the vibration and the interior sound pressure of the shell by employing the techniques of subsystem synthesis and modal expansion where the external distributed pressures are chosen according to experimental data of an actual aircraft. The absorbers are successful for attenuating the vibration and the noise. The effects of altering various parameters of the shell, the external pressures, and the absorbers, are also studied and discussed. Finally, some general guidelines of absorber design for vibration and noise control of the fuselage are presented.

Abstract: A comprehensive analytical methodology has been developed for predicting the onset of widespread fatigue damage (WFD) in fuselage structure. The determination of the number of e ights and operational hours of aircraft service life that are related to the onset of WFD includes analyses for crack initiation, fatigue crack growth, and residual strength. Therefore, the computational capability required to predict analytically the onset of WFD must be able to represent a wide range of crack sizes, from the material (microscale) level to the global (structural-scale ) level. The results of carefully conducted teardown examinations of aircraft components indicate that fatigue crack behavior can be represented conveniently by the following three analysis scales: 1 ) small three-dimensional cracks at the microscale level, 2 ) through-the-thickness two-dimensional cracks at the local structural level, and 3 ) long cracks at the global structural level. The computational requirements for each of these three analysis scales are described in this paper.

Abstract: An aircraft passenger extraction system having a fuselage and a tail section removably attached to the fuselage. A plurality of interconnected passenger modules are removably disposed within the fuselage. Four I-beam rails are longitudinally mounted about the inside circumference of the fuselage, and the passenger modules have four sets of wheels that slide on the rails. The tail section has four pairs of separation flaps and four extraction flaps to assist in the separation of the tail section in the event of an airborne emergency. Upon separation of the tail section of the fuselage, the interconnected passenger modules are slidingly withdrawn from the fuselage. Once the passenger modules clear the fuselage, the interconnections of each passenger module are severed in sequence, and a series of parachutes are deployed to safely float each passenger compartment to earth.

Abstract: The resulta of residual strength pressure tests and nonlinear analyses of stringer-and frame-stiffeded aluminum fuselage panels with longitudinal cracks are presented. Two types of damage are considered: a longitudinal crack located midway between stringers, and a longitudinal crack ajdacent to a stringer and along a row of fasteners in a lap joint that has multiple-site damage (MSD). In both cases, the longitudinal crack is centered on a severed frame. The panels are subjected to internal pressure plus axial tension loads. The axial tension loads are equivalent to a bulkhead pressure load. Nonlinear elastic-plastic residual strength analyses of the fuselage panels are conducted using a finite element program and the crack-tip-opening-angle (CTOA) fracture criterion. Predicted crack growth and residual strength results from nonlinear analyses of the stiffened fuselage panels are compared with experimental measurements and observations. Both the test and analysis results indicate that the presence of MSD affects crack growth stability and reduces the residual strength of stiffened fuselage shells with long cracks.

Abstract: Many traditional data bases, which involved smooth-sided forebodies, are no longer relevant for designing advanced aircraft. The current work provides data on the impact of chined-shaped fuselage cross section on the stability of a generic fighter configuration. Two different chined-shaped fuselages were tested upright and inverted. It was found that a fuselage with a 30 degree included chine angle resulted in significantly higher values of Cl,max than a fuselage with a 100 degree included chine angle. This difference was attributed to the more beneficial vortical interaction between the stronger forebody vortices coming off of the sharper chine edges and the wing vortices. The longitudinal stability of the configuration with the sharper chine angle was also better because, based on pressures and flow visualization, the vortex burst over the wing was delayed until significantly higher values of alpha. Unstable rolling moment derivatives were also delayed to higher values of alpha for the sharper chine angle cross section. Furthermore, it was found that directional stability of both of the upright configurations, which had larger lofts in cross section above the chine lines than below the chine lines, was better than for the inverted configurations.

Abstract: The invention is related to air flight vehicles, such as vertical take-off and landing (VTOL) airplanes, helicopters and covercraft. The goal of this invention is to create an air rotor designed so that while after vertical take off or cover regime, one can be stopped, fixed in a specific position and hidden into the fuselage (gondola) thus eliminating of air resistance when the rotor is not in working state. On landing this rotor can be extended out, brought into rotation and used for creation of lift force and vertical landing. The indicated goal is achieved by means of the rotor made as single blade (uniblade). The author solved the problem of force and moment balance of single blade. The center of gravity of the counterweight is located below the horizontal plane, and the blade has the horizontal sway axle, that crosses the vertical rotor rotation axis. The author offer this rotor on single axis, on co-axis, and on different exiles. This uniblade rotor is designed to subsonic and supersonic VTOL airplanes, for helicopter cars, flight motorcycles, hoppycopters, and covercraft. The uniblades be used also as a veritable sweep wing (for supsonic and supersonic aircraft).

Abstract: An aerial decoy comprising a fuselage having forward and aft ends. Disposed within the fuselage are a plurality of decoy discs. Rotatably connected to the forward end of the fuselage is a ram air turbine which is cooperatively engaged to the decoy discs such that the rotation of the ram air turbine facilitates the dispensation of the decoy discs from the aft end of the fuselage.

Abstract: A comparison of active smart structure - piezoelectric control systems and aerodynamic active systems for vibration alleviation and elastic mode damping of a military aircraft structure is presented. The vibration alleviation systems which are operative at flight in turbulence or during maneuvers at high incidence corresponding to severe buffeting conditions are under investigation by DASA as a part of research study on advanced aircraft structures. The active systems for elastic mode damping (ASD) are designed as digital systems to provide vibration alleviation and have an interface to the Flight Control System (FCS) or are directly part of the FCS. The sensor concept of all different systems is the same as the sensor concept used for the FCS with the corresponding benefits of redundancy and safety. The design of systems and the comparisons of system properties are based on open and closed loop response calculations, performed with the dynamic model of the total aircraft including coupling of flight mechanics, structural dynamics, FCS dynamics and hydraulic actuator or piezo-actuator dynamics. Aerodynamic systems, like active foreplane and flap concepts, rudder and auxiliary rudder concepts, and piezoelectric systems, like piezo interface at the interconnection fin to rear fuselage and integrated piezo concepts are compared. Besides the essential effects on flexible aircraft mode stability and vibration alleviation factors system complexity and safety aspects are described.

Abstract: An assembly cell for accurate placement of stringer clips in a channel of a stringer for a compound contour section of an airplane fuselage, and for holding the clips at a predetermined orientation while drilling holes for fastening the clips in the stringers includes a fixture having locating surfaces for holding the stringer in a configuration that is the same as the configuration it will have when installed at its designated position in the airplane. The fixture has headers adjustable to conform to the shape of the airplane in which the stringers will be installed, The headers have clamps for securing the stringers against accurately machined reference surfaces on the headers. An end effector is held by a machine tool in the assembly cell for gripping a stringer clip and inserting it into the stringer channel. The end effector has spreader bars for insertion in the stringer channel operable to spread sidewalls of the stringer and widen the channel to facilitate insertion of the stringer clip, and has clamp bars positionable on opposide sides of the stringer and operable to close on the stringer to squeeze the stringer sidewalls on the stringer clip after insertion in the channel. Opposed drills on the end effector have right angle drives holding drill bits for simultaneously match drilling holes through the sidewalls and through the stringer clip. The end effector machine tool operates as a positioning mechanism for accurately for positioning the drills at each stringer clip location along the stringer clip to positions corresponding to the contour of the fuselage at each of the locations under control of a CNC machine controller.

Abstract: Vertical lift in an aircraft is produced by driving a column of air downwardly, through an annular thrust-flow channel which is formed in the body (fuselage) of the aircraft. The aircraft also has an aerodynamic shape which is capable of developing lift responsive to forward flight. The annular thrust-flow channel is provided with a flow control mechanism which is capable of directing the developed air flow in varying orientations between a substantially vertical (axial) orientation for developing stationary, vertical lift (i.e., hovering) and a vectored (angled) orientation for developing a vertical component for producing lift and a horizontal component for producing forward (or rearward) flight, or flight to either side.

Abstract: An improved version of the Wall Signature Method was developed to compute wall interference effects in three-dimensional subsonic wind tunnel testing of aircraft models in real-time. The method may be applied to a full-span or a semispan model. A simplified singularity representation of the aircraft model is used. Fuselage, support system, propulsion simulator, and separation wake volume blockage effects are represented by point sources and sinks. Lifting effects are represented by semi-infinite line doublets. The singularity representation of the test article is combined with the measurement of wind tunnel test reference conditions, wall pressure, lift force, thrust force, pitching moment, rolling moment, and pre-computed solutions of the subsonic potential equation to determine first order wall interference corrections. Second order wall interference corrections for pitching and rolling moment coefficient are also determined. A new procedure is presented that estimates a rolling moment coefficient correction for wings with non-symmetric lift distribution. Experimental data obtained during the calibration of the Ames Bipod model support system and during tests of two semispan models mounted on an image plane in the NASA Ames 12 ft. Pressure Wind Tunnel are used to demonstrate the application of the wall interference correction method.

Abstract: Deals with the configuration of autonomous underwater vehicle (AUV) for inspection of underwater cables considering low hydrodynamic drag, sensor alignment, collision avoidance manoeuvre, turning manoeuvre and rolling motion. The AUV consists of a fuselage of body of revolution with low hydrodynamic drag, fore and aft horizontal wings, upper and lower vertical tails and a pair of horizontal thrusters at both sides of the aft horizontal wings. The shape of the forward horizontal wing and that of the rear horizontal wing can be determined by use of a nonlinear optimization method under the constraints of magnetic sensor alignment for cable tracking, dynamic stability in the vertical plane and the performance of the collision avoidance manoeuvre. The performance of the turning manoeuvre and rolling motion are attributed to thrust force difference between a pair of thrusters at both sides of the rear horizontal wing, because the shapes of the upper and lower vertical tails can be designed from the viewpoint of the dynamic stability in horizontal plane.

Abstract: In accordance with the present invention, there is provided a variable camber delta-shaped aircraft. The aircraft is provided with an integrated fuselage/wing generally defining the aircraft and having longitudinal and lateral axes. The fuselage/wing has a forward section which is rotably attached to an aft section about the lateral axis. The aircraft is further provided with an aerodynamic lifting surface which is disposed about the fuselage/wing and defined by a camber. The forward section has a downwardly deflected position when rotated relative to the aft section. The forward and aft sections are cooperatively formed to increase the camber of the lifting surface when the forward section is in the deflected position.

Abstract: An aircraft engine is attached to the fuselage by a mounting structure that has a vibration absorbing system which includes first and second sensors to produce signals indicating vibration along two orthogonal axes. A third sensor, such as an accelerometer, is coupled to the mounting structure for sensing vibration of the engine. A tachometer circuit, connected to the third sensor, produces first and second speed signals that indicate the speeds of two rotating spools in the engine. Four vibration absorbers are attached to the engine mounting structure and controllers are provided to dynamically tune the resonant frequency of each vibration absorber in response to a unique combination of one of the two vibration signals and one of the two speed signals. Thus the vibration absorbers reduce the engine spool vibrations that are transmitted through the mounting structure along the two orthogonal axes. The resonant frequency of each absorber is altered to track variation of the vibrations due to changes in engine speed.

Abstract: Rotor ine ow aerodynamics and noise measurements were made on three cone gurations of an 8% scale model of the XV-15 tiltrotor in hover: 1 ) single isolated rotor, 2 ) two rotors with no fuselage, and 3 ) complete tiltrotor aircraft. For the tiltrotor aircraft cone guration and for the cone guration of two rotors without the fuselage, the mean ine ow velocity was higher at c = 270 deg compared with the rest of the rotor disk, leading to reduced blade angle of attack and blade loading in this region. This azimuthally varying blade loading caused an impulsive noise that radiated preferentially behind the model. For the complete tiltrotor cone guration, the turbulence ingested by the rotors was intermittent and depended on the instantaneous position of the fountain e ow, which shifted from side to side across the longitudinal plane of the model. The fountain turbulence had a higher velocity scale, smaller length scale, and was closer to isotropic than the ingested ambient turbulence. The tiltrotor cone guration radiated less harmonic noise, but more broadband noise than the cone guration with two rotors and no fuselage. Diagonal fences on the wings of the tiltrotor reduced the ine ow turbulence intensity in the fountain region by a factor of about 3, and reduced the noise by 4.1 dBA behind the model. Scaling relations were derived to extrapolate the model measurements to the full-scale XV-15.

Abstract: An aircraft with a variable forward-sweep wing and the method of configuring the wing in an optimal position for a desired flight regime. The variable forward-sweep wing is positionable from an essentially unswept position to a full-forward sweep position. In the unswept position the wing is approximately orthogonal to a fuselage centerline, while in the full-forward sweep position the wing has approximately a delta wing planform. Moreover, as the wing position changes from the unswept position to the full-forward sweep position the trailing edge becomes the leading edge. In addition, the aforementioned apparatus may be used in a method to configure the aircraft for flight in a desired flight regime. This method includes moving the wing to an optimal position for the desired flight regime.

Abstract: : A commuter category Beechcraft 1900C airliner was subjected to a vertical impact drop test at the FAA William J. Hughes Technical Center, Atlantic City International Airport, New Jersey. The purpose of this test was to measure the impact response of the fuselage, cabin floor, cabin furnishings (including standard and modified seats), and anthropomorphic test dummies. The test was conducted to simulate the vertical velocity component of a severe but survivable crash impact. A low-wing, 19-passenger fuselage was dropped from a height of 11' 2" resulting in a vertical impact velocity of 26.8 ft/sec. The airframe was configured to simulate a typical flight condition, including seats (normal and experimental), simulated occupants, and cargo. For the test the wings were removed; the vertical and horizontal stabilizers were removed; the landing gear was removed; and the pilot and copilot seats were not installed. The data collected in the test and future tests will supplement the existing basis for improved seat and restraint systems for commuter category 14 Code of Federal Regulation (CFR) Part 23 airplanes. The test article was fully instrumented with accelerometers and load cells. Seventy-nine data channels were recorded. Results of the test are as follows: - the fuselage experienced an impact in the range of 149-160 g's, with an impact pulse duration of 9-10 milliseconds - the simulated occupants experienced g levels in the range of 32-45 g's with a pulse duration of 44-61 milliseconds - the test was considered to be a severe but definitely survivable impact - the fuselage structure maintained a habitable environment during and after the impact - the seat tracks remained attached to the fuselage along the entire length of the fuselage - all standard seats remained in their tracks after the impact - all exits remained operable