Abstract: The thin-walled structures of interest in this chapter are generally enclosed regions, as typified by an aircraft fuselage. They are modeled as combinations of folded plates and cylindrical shell segments as shown in Figure 8.1. Other examples of structures which fall into this category include plates with stringers, corrugations, channels, ducts, troughs, open and closed thin-walled tubes with one or more cells.

Abstract: An aircraft adapted for flight in helicopter mode with its longitudinal axis oriented generally vertically and in airplane mode with its longitudinal axis oriented generally horizontally is provided with the capability of launching and landing with the tail end directed skyward. The invention also includes improvements to the controllability and efficiency of aircraft in helicopter mode provided by the stabilizer wings and relative rotation of the fuselage section about the aircraft's longitudinal axis. The aircraft has an elongate boom positioned between the rear fuselage and the stabilizer wings for engagement with a base structure. The base structure may be attached to a building, a trailer transporter, a ship or some other structure. The base may be a beam having latching arms that swing inward when pressure from the boom is applied, then trap the boom in a recess between the latching arm and beam, thereby suspending the aircraft. An actuator for releasing the latching arms is provided to allow the aircraft to launch.

Abstract: A method and apparatus for controlling the flight direction of a blimp. The blimp includes an elongate fuselage for containing a lighter-than-air gas, a movable rudder located toward a tail end of the fuselage, and at least one propeller mounted on a side of the fuselage. The propeller is attached to the blimp by a connection that provides two axes of movement. The propeller connection can be rotated or pivoted (1) about a first axis which is a horizontal axis that is perpendicular to a direction of the longitudinal axis of the elongate fuselage, and/or (2) about a second axis which is a vertical axis that is perpendicular to the direction of the longitudinal axis of the fuselage.

Abstract: An apparatus and method for forming an aircraft fuselage having at least one doubly curved portion. A plurality of fuselage panels are extruded with integrally formed stiffening portions extending preferably longitudinally the full length of each fuselage panel. The stiffening portions are further formed so as to be generally parallel to one another. The fuselage panels are then formed into skin panels having a desired curvature and trimmed into gores, and the stiffening portions trimmed at end portions that run out toward longitudinal edges of the panels. When adjacent panels are affixed together, the stiffening portions run out toward longitudinal edges of adjacent panels. The longitudinal edges are secured by splice longerons which are riveted or welded to the longitudinal edges of the skin panels. Independent fastening members are then secured to the stiffening portions and the splice longerons at the points where the stiffening portions run out to positions closely adjacent the splice longerons. The apparatus and method permits aircraft fuselages having doubly curved portions to be formed from extruded panels which is more cost efficient and makes use of a greater percentage of raw materials than conventional methods of machining such panels out of non-near-net stock such as plate.

Abstract: In aircraft, much of the noise in the cabin is a result of pressure fluctuations on the exterior fuselage skin, which vibrates the skin and causes sound pressure waves to propagate through the trim panel cavity, exciting the trim panels, which radiate into the cabin. In the present invention, noise is reduced in the cabin of an aircraft by reducing the acoustical impedance of the trim panel cavity of the aircraft using an array of independent active sound absorbing acoustic cells. Acoustical drivers are mounted within the trim panel cavity of the aircraft. Each acoustical driver has at least one electroacoustical sensor mounted in its immediate vicinity. The sensor is connected to the driver to provide negative feedback, which has the effect of canceling noise in the vicinity of each cell. Since the effect of each cell is to absorb acoustical energy, the acoustical impedance of the trim panel cavity is reduced. Multiple sound absorbing acoustic cells in the trim panel cavity therefore decrease the acoustical impedance inside the trim panel cavity which in turn reduces the transmission of noise into the cabin.

Abstract: A mathematical model and a computer code have been developed to fully couple the vibration of an aircraft fuselage panel to the surrounding flow field, turbulent boundary layer and acoustic fluid. The turbulent boundary layer model is derived using a triple decomposition of the flow variables and applying a conditional averaging to the resulting equations. Linearized panel and acoustic equations are used. Results from this model are in good agreement with existing experimental and numerical data. It is shown that in the supersonic regime, full coupling of the flexible panel leads to lower response and radiation from the panel. This is believed to be due to an increase in acoustic damping on the panel in this regime. Increasing the Mach number increases the acoustic damping, which is in agreement with earlier work.

Abstract: A fixed-wing four-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 short fuselage for convenient roadability and garageability, with horizontal stabilizer of significant span with foldable tips for adequate flight stability. 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 adequate ground handling. Ground propulsion is by automotive-style transmission driving the front wheels.

Abstract: An air-chute safety system (10) for an aircraft (12) having a fuselage (14) with a cockpit (16) and a tail end (18), and a pair of wings (20) with engines (22). The system (10) comprises a plurality of drag parachutes (24) stored in and deployed from the tail end (18) of the fuselage (14), in conjunction with an aerodynamic braking apparatus in the aircraft (12), which will shut down the engines (22) during an emergency when in flight, so that the aircraft (12) will slow down. A plurality of top parachutes (26) are provided. A structure (28) is for storing each of the top parachutes (26) in vertical spaced apart relationships below a top surface (30) of the fuselage (14) between the cockpit (16) and tail end (18) thereof. A component (32) is for deploying each of the top parachutes (26) from above a bottom surface (34) of the fuselage (14) and below each of the storing structures (28), through the top surface (30) of the fuselage (14), after the drag parachutes (24) are deployed, so that the aircraft (12) will slowly and safely float down to the earth.

Abstract: Conventional ultrasonic or eddy current inspection of structures requires a probe to be scanned over the whole area to be tested. This is extremely time consuming, and hence costly, when large areas such as aircraft wings or pressure vessels are to be covered. A further disadvantage of ultrasonic inspection is that a coupling fluid between the transducer and the structure is generally required. The most reliable coupling method is to use immersion, the testpiece being fully immersed in a water bath. However, with large structures this is frequently not practical and they are often tested using jet probe assemblies, the ultrasound being propagated down jets of water directed at the structure. However, this is generally only practical at the manufacture stage and field inspection is often carried out manually using contact transducers, coupling being achieved by applying gel to the surface of the structure. Many structures also have critical areas which are difficult to inspect because access is impeded. For example, spars and other stiffeners in aircraft pose problems because once the aircraft is built they are covered by the fuselage or wing skin.

Abstract: A wall element has a noise attenuating characteristic and is suitable for use as an aircraft cabin interior panel element that is mounted directly on the aircraft fuselage structure. The wall element includes a rigid lightweight composite panel having a core arranged between two cover layers, and an outer layer arranged adjacent and spaced away from one of the cover layers. The core and the cover layers of the composite panel are air permeable in a direction through the thickness of the panel, while the outer layer is non-air-permeable and is relatively soft and flexible. The outer layer is connected to the composite panel preferably only along the perimeter edge thereof. The non-permeable outer layer faces the aircraft cabin space, while the permeable composite panel is mounted on the fuselage structure. With this structure, acoustic vibrations of the fuselage, to which the wall element is attached, are attenuated and do not result in substantial noise radiation from the outer layer into the cabin.

Abstract: During the NASA sponsored Advanced Technology Composite Aircraft Structures (ATCAS) research program conducted at Boeing, a composite cocured skin/stringer/frame construction was identified as a potentially cost effective structural configuration for the fuselage of a wide-body civilian transport aircraft. One of the critical areas of the design was identified as the bonded interface between the skin and frame. The outline of a global/local approach to analyzing the bondline under internal pressure and axial fuselage loading is presented. Results from a geometrically nonlinear finite element (FE) of a panel bay are combined with the results of detailed local FE models of the frame to skin bondline. Both a strength of material and a strain energy release rate approaches are used to examine the onset of failure and its subsequent propagation.

Abstract: An aircraft pressure containment module is adapted to be mounted generally transversely within an aircraft fuselage for providing pressure containment to separate pressurized and non-pressurized sections of an aircraft. A plurality of web segments are joined together to form a web member. A ring member is arranged on a first pressurized side of the web member. A plurality of truss assemblies are also arranged on the first side of the web member. Each truss assembly is comprised of a first truss member adapted to extend outwardly from the ring member to a peripheral edge of the web member. The first truss member is joined to the web member and joins together adjacent web segments. A second truss member extends from the first truss member in a direction generally normal to the first side of the web member. The second truss member is adapted to be joined to an aircraft fuselage in a longitudinal sense or it may comprise a portion of longitudinally extending member of the fuselage such as a stringer. A third truss member is joined to the second truss member and the first truss member to form a truss assembly having at least in part a generally triangular shape. The truss assemblies are adapted to react pressure loading from the web member to the fuselage of an aircraft in a manner similar to a conventional dome type pressure containment module. An aircraft using the module has an elongated fuselage extending in a longitudinal direction and the pressure containment module mounted generally transversely within the fuselage.

Abstract: A fixed wing aircraft that can be automatically and rapidly converted to a roadway vehicle within seconds while driving in the highway, comprising a one-piece wing rotatably mounted on top of the fuselage The aircraft features a 3-horizontal-surface design allowing good pitch stability and damping in a very short fuselage, stall-resistant feature and reduction in induced drag at cruise speed The aircraft's wing is enlargeable with additional wing tip segments allowing boost in aspect ratio hence improving efficiency at high loads The vehicle has low center of gravity, four wheels with independent suspension, nose-height leveling for take-off and landing, and anti-sway mechanism in order to prevent roll-over in a tight turn Ground propulsion is by both a ducted fan and hydraulic rear wheel drive allowing rapid accleration in the roadable mode

Abstract: An investigation has been conducted in the NASA Langley Research Center’ s 14 by 22 ft Subsonic Tunnel to further the development of semispan testing capabilities. A twin-engine, energy efe cient transport model with a four-element wing in a takeoff cone guration was used for this investigation. Initially, a full-span cone guration was tested, and force and moment data and wing and fuselage surface pressure data were obtained as a baseline data set. The semispan cone gurations were then mounted on the windtunnel e oor, and the effects of fuselage standoff height and shape were investigated. Results indicate that the semispan cone guration was sensitive to variations in standoff height, and that a standoff height equivalent to 30% of the fuselage radius resulted in better correlation with full-span data than no standoff or the larger standoff cone gurations investigated. Undercut standoff leading edges improved the correlation of semispan data with full span data in the region of maximum lift coefe cient.

Abstract: A pressure bulkhead especially for closing the tail end of the pressurized interior of an aircraft fuselage includes a curved or cupped disk, a peripheral rim, and a transition portion joining the rim to the disk. All the components of the bulkhead are integrally formed of a fiber-reinforced composite material. The peripheral rim and the transition portion together form a circumferential frame that supports circumferential forces and transfers forces into the aircraft fuselage, to which the rim is connected. The transition portion has a curved cross-section that curves in the opposite direction relative to the curvature of the cupped disk, whereby the rim bends or flares back toward the convex side of the cupped disk. The curvature of the transition portion is relatively tight so that an acute angle is formed between the rim and the cupped disk, and the rim extends in contact with the inner surface of the fuselage structure over a contact range at which the rim may be rivet-connected to the fuselage. A reinforcing ring may be arranged between the rim and the fuselage.

Abstract: The problem of rotor-fuselage aerodynamic interaction has to be considered in industry applications from various aspects. First, in order to increase helicopter speed and reduce operational costs, rotorcraft tend to be more and more compact, with a main rotor closer to the fuselage surface. This creates significant perturbations both on the main rotor and on the fuselage, including steady and unsteady effects due to blade and wake passage and perturbed inflow at the rotor disk. Furthermore,the main rotor wake affects the tail boom, empennage and anti-torque system. This has important consequences for helicopter control and vibrations at low speeds and also on tail rotor acoustics (main rotor wake-tail rotor interactions). This report describes the US Army-France MOD cooperative work on this problem from both the theoretical and experimental aspects. Using experimental 3D velocity field and fuselage surface pressure measurements, three codes that model the interactions of a helicopter rotor with a fuselage are compared. These comparisons demonstrate some of the strengths and weaknesses of current models for the combined rotor-fuselage analysis.

Abstract: A technique for augmenting the aileron to increase the roll rates of a high-performance aircraft at high dynamic pressures for enhancement of the rolling maneuver is examined. Antisymmetric twist and camber distribution on a realistic flexible wing is determined to counteract the detrimental twisting moment of the aileron rotation to achieve recovery of the lost roll rates at high dynamic pressures. A method for prescribing the antisymmetric wing twist and camber distribution by reversing the twist and camber resulting from use of the aileron only is described. The retwisting and recambering of the wing is achieved by providing control forces obtained from a technique referred to as fictitious control surfaces. The technique of retwisting and recambering a flexible wing demonstrates a full recovery of roll rate at all dynamic pressures. Here, a full-scale realistic wing is considered for the assessment of strain energy as a measure of necessary control energy required to produce the antisymmetric twist and camber deformation on the aileron-twisted wing to recover the lost roll performance.

Abstract: The Advanced Technology Composite Aircraft Structures (ATCAS) program has studied transport fuselage structure with a large potential reduction in the total direct operating costs for wide-body commercial transports. The baseline fuselage section was divided into four 'quadrants', crown, keel, and sides, gaining the manufacturing cost advantage possible with larger panels. Key processes found to have savings potential include (1) skins laminated by automatic fiber placement, (2) braided frames using resin transfer molding, and (3) panel bond technology that minimized mechanical fastening. The cost and weight of the baseline fuselage barrel was updated to complete Phase B of the program. An assessment of the former, which included labor, material, and tooling costs, was performed with the help of design cost models. Crown, keel, and side quadrant cost distributions illustrate the importance of panel design configuration, area, and other structural details. Composite sandwich panel designs were found to have the greatest cost savings potential for most quadrants. Key technical findings are summarized as an introduction to the other contractor reports documenting Phase A and B work completed in functional areas. The current program status in resolving critical technical issues is also highlighted.

Abstract: A secondary wing system for use on an aircraft augments the lift, stability, and control of the aircraft at subsonic speeds. The secondary wing system includes a mechanism that allows the canard to be retracted within the contour of the aircraft fuselage from an operational position to a stowed position. The top surface of the canard is exposed to air flow in the stowed position, and is contoured to integrate aerodynamically and smoothly within the contour of the fuselage when the canard is retracted for high speed flight. The bottom portion of the canard is substantially flat for rotation into a storage recess within the fuselage. The single canard rotates about a vertical axis at its spanwise midpoint. The canard can be positioned between a range of sweep angles during flight and a stowed position in which its span is substantially parallel to the aircraft fuselage. The canard can be deployed and retracted during flight. The deployment mechanism includes a circular mounting ring and drive mechanism that connects the canard with the fuselage and permits it to rotate and to change incidence. The deployment mechanism further includes retractable fairings which serve to streamline the wing when it is retracted into the top of the fuselage.

Abstract: The aircraft consists of a fuselage (11) with a wing (19) mounted on fins above its rear end, a horizontal stabiliser (40) on its nose, and two arms (27,29) projecting from opposite sides of the middle of the fuselage It has airscrews (37) mounted on nacelles (33,35) pivoted to the ends of the arms so they can be rotated between horizontal and vertical positions The airscrews are driven by a single power unit (17) in the rear of the fuselage through a transmission system (45) The aircraft's centre of gravity is forward of the side arms, and the tips of its wind have downward pointing fixed stabiliser fins The nose stabiliser consists of a horizontal fin (41) on each side of the nose, equipped with an variable flap (43) The airscrew transmission system comprises shaft which passes through the side arms and is connected to the variable-pitch screws through bevel gear mechanism

Abstract: An insulation system for the fuselage walls of an aircraft. The insulation is a formed-in-place, sprayed-on foam that is applied against a barrier material that is draped against the inside of the exterior skin of the fuselage and over any inwardly projecting structural members attached to the exterior skin. No substantial portion of the barrier material passes underneath any laterally projecting portions of the structural members, so that the subsequently solidified, semi-flexible foam (i.e. insulation) can be removed at a later time for necessary inspections of the aircraft, without being trapped behind these lateral projections. The foam is preferably fire retardant and hydrophobic. An additional protective layer may also be sprayed on to the outside surface of the insulating foam.

Abstract: A rigid cargo barrier for a cargo airplane includes a lightweight structural bulkhead that is attached to the floor and fuselage shell of the aircraft and a plurality of compact attachments that connect the bulkhead to the floor and fuselage shell of the aircraft. The rigid cargo barrier also includes a restraint system that connects the bulkhead to the aircraft floor and fuselage shell. The restraint system restrains fuselage expansion under crash loads to maintain the bulkhead in place during a crash event and, hence, provide protection to the crew from shifting cargo.

Abstract: A wing structure for hang gliders, ultralites, gliders, heavy aircraft, ornithopters and sailsoaring flying boats which has a first wing set that pivots on two axis, and a second wing set that remains substantially immobile relative to the fuselage or keel. The wing pivots on the lateral axis of the fuselage or keel by moving along a slider assembly that allows it to move from a swept position to a more perpendicular position relative to the keel to control the amount of lift. The wing also pivots on the longitudinal axis of the fuselage or keel to control banking. On hang glider versions of the device, optional ducted fan, propeller, or jet propulsion units provide power to maintain flight. An optional bungee launch assembly assists takeoff from relatively flat surfaces. An optional "telepresence" wing controller allows the pilot to control wing motion with minimal physical exertion. Optional landing gear are controlled by wing position, such that the landing gear are lowered when the wings are in the forward takeoff/landing position and raised when the wings are in the swept flight position.

Abstract: Applying system identification methodology, an aerodynamic database covering the entire operational flight envelope, was estimated from Transall flight data. We emphasize the modeling and identification of some of the Transall specific operational characteristics such as the ramp door and dropping of heavy loads. In addition, aerodynamic effects because of speed brakes and landing gear are determined applying a nonlinear maximum likelihood parameter estimation method. The identification results indicate that the speed brakes function as pure drag generators only for the retracted landing flaps. For extended landing flaps, applying speed brakes affects the aircraft pitching motion. The landing gear is found to affect both the longitudinal and lateral-directional aircraft motion. Opening of the ramp door results in an effectively enlarged fuselage, affecting mostly the lateral-directional motion. It is demonstrated that the load dropping can be adequately modeled by properly accounting for the variations in the aircraft mass characteristics.

Abstract: A ground floodlight for an aircraft having a fuselage including a forward cargo door includes a generally flat lens having a first region for directing light toward the front cargo door and a second region for directing light generally forward to illuminate ground in front of the aircraft. The lens is generally flat and smooth and is adapted for flush mounting in the fuselage. A corresponding method for improving ground floodlight illumination includes directing light generated from a lamp within the fuselage in an outward direction and then redirecting first and second portions of the light toward the front cargo door and the ground in front of the aircraft, respectively.

Abstract: A fuselage door (2) for a pressurized aircraft fuselage (1) in which the hoop tension loads caused by cabin pressure are carried through the door rather than around the door. Stationary interlocking moldings (104) in the door (2) and fuselage door (32, 39) seat transfer loads through the door-fuselage intersection. The interlocking portions of the moldings (104) have a sloped engaging portion (6) having an angle (B) such that the inward directed force component of tension due to a pressure differential is less than the outward directed force from cabin pressure on the door.

Abstract: A new business jet is optimized to fly significantly faster than most current production aircraft while operating from relatively short runways. This new airplane is required to accommodate eight passengers in a double-club arrangement and to carry six passengers for 2800 n mile at a Mach number between 0.81-0.85. Two aircraft optimization codes are used here to ensure the validity of the design results and to identify errors in the analysis methods. These codes include the aircraft analysis methods necessary to evaluate the aircraft performance over an entire mission and optimization routines that enable the development of a family of optimum configurations. The design objective, empty weight, is shown to change approximately 1% between 30-40 deg of wing sweep at a Mach number of 0.81. At a fixed wing sweep of 31.5 deg and a Mach number of 0.81, the empty weight decreases less than 3.5% when the wing's thickness-to-chord ratio is increased from 0.10 to 0.14. A study of the design's sensitivity to Mach number indicated that the optimum empty weight and wing thickness began to change rapidly between the Mach numbers of 0.83-0.85.