Abstract: Design studies have been performed on continuous filament wound grid-stiffened composite structural concepts for aircraft fuselages. Stability equations based on a Donnell-type theory, which includes transverse shear deformation in the unidirectional composite ribs, have been developed for a general grid-stiffened circular cylindrical shell. Three candidate design concepts—isogrid, orthogrid, and generalized orthogrid were compared. The loading cases considered were uniaxial compression and combined axial compression and torsion. All three concepts are weight competitive. The isogrid concept, however, is the most attractive due to its demonstrated damage tolerance characteristics.

Abstract: A high altitude, ultra lightweight aircraft has broad area relatively flat wings, and is provided with solar cells on its upper wing surfaces, and long wavelength, infrared responsive cells on its lower wing and fuselage surfaces, to provide enough power to maintain the aircraft aloft at high altitudes indefinitely. The aircraft is intended to fly at altitudes of about 60,000 feet to 110,000 feet, at speeds in the vicinity of 70 to 200 miles per hour. A battery or fuel cell and electrical system is provided to store excess electrical power and to supply electrical power when needed. A large, low speed, pusher propellor is driven by an electrical motor. The aircraft may be launched with an opaque balloon, which not only raises the aircraft to the desired elevation but also conceals it.

Abstract: It is well established that there is a strong coupling between body motion and boundary layer separation with attendant vortex shedding. In the present paper this coupling is studied for the particular case of a missile or an aircraft fuselage at very high angles of attack. It is shown that the unusual results obtained in recent tests can be explained by considering the so-called "moving-wall effect" on boundary layer transition and/or separation.

Abstract: A model aircraft kit which enables the purchaser to assemble a variety of model aircraft configurations and planforms, uses a relatively small number of components. The components of the model aircraft kit comprise at least one of several possible variations of fuselages, a wing panel or panels, and a number of stabilizing and control surfaces which may include conventional or diagonally hinged control surfaces. Releasable connectors which may include a pair of releasable wing tip appendage connectors, are used to join the components thereby providing a wide variety of planforms and configurations. At least one of the fuselages is capable of being separated into subparts one of which becoming a short fuselage being capable of accommodating some or all of the remaining components of the model aircraft kit. Several individually unique model aircraft are assembled using a pair of wing tip appendage connectors with respecting stabilizing surfaces, independent of other multiple configuration features of the aircraft kit.

Abstract: A mechanism and method for folding the wing of a tilt-rotor aircraft, hav a rotatable load bearing member cooperatively disposed between a high point of the fuselage and the tilt-rotor wing, and a plurality of locking, load bearing pin joints disposed between the fuselage and the tilt-rotor wing, concentric to the rotatable load bearing member, and wherein the tilt-rotor wing may be stored by rotating the tilt-rotor wing, as a single segment, about the rotatable load bearing member, to a position substantially parallel to the fuselage, and wherein the wing may be locked in the flight configuration and the storage configuration with the same pin joints.

Abstract: In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Abstract: The present invention provides a single seat aircraft having the capabilities of vertical takeoff, landing and hovering operations utilizing the X-Wing as a conventional helicopter rotating wing. After transition to forward flight following takeoff, the rotating wing is stopped and becomes a fixed wing of "X" configuration. The aircraft utilizes two engines within the fuselage, one engine being positioned vertically above the other along the longitudinal axis of the aircraft, the particular arrangement of engines allowing aircraft size and weight to be substantially reduced.

Abstract: Earlier work that produced generalized rules for obtaining trajectories of water droplets moving around aircraft fuselages has been extended to cover the case for particles of arbitrary shape. The parameters determining the major features of the trajectories are the fuselage radius and a generalized Stokes number that can be used to calculate an equivalent water drop size for particles of arbitrary shape and density. As for water drops, the maximum width of the shadow zone for sampling the particles is about one-fifth of the fuselage radius, and it occurs for those in columns and plates of typical densities whose major dimensions are about 600 μm and 2 mm, respectively.

Abstract: An amphibious executive or business airplane, also having military application, possesses sponsons integrated with the airplane fuselage. The trailing end of the sponsons end in a sharp edge, thus eliminating the need for a hull or float step or venting. A channel is formed by the sponsons and the underside of the fuselage, which channel has a constant cross-section and produces a surface lift effect facilitating take-off and landing. A pair of keels beginning substantially aft of the center of gravity of the airplane provide directional stability on the water. The aircraft is preferably jet powered with the inlet of the engines disposed above the sponsons to eliminate water intake. Variable geometry wings of telescoping design maximize aerodynamic performance on take-off and landing and during the high altitude, high speed cruise regime of the flight.

Abstract: Design development tests were conducted to investigate the crashworthy characteristics of composite helicopter fuselage subcomponents, and to design helicopter center beam/bulkhead specimens lighter than structural elements of honeycomb sandwich construction. Skinstringer designs of center beams - made of carbon, and hybrids of carbon and Kevlar - were fabricated and tested in axial compression. Crashworthy design parameters of specific energy, operating load and stroke efficiency were investigated. 8 references, 15 figures, 2 tables.

Abstract: A fail-safe, automatic interlock connector arrangement automatically electrically interconnects an electrical component on a wing to another electrical component on a fuselage of a radio-controlled model airplane at the same time that the wing is connected to the fuselage.

Abstract: Interior sound in aircraft fuselages is directly coupled to fuselage vibration. This suggests the use of active vibration control of the fuselage vibration to reduce interior noise levels. In this paper, preliminary experiments to investigate this new technique are discussed. The experiments were performed on a closed cylindrical shell excited by an external acoustic source, representative of an aircraft fuselage. Active vibration control of the shell was achieved by a single minishaker attached pointwise to the exterior wall of the cylinder. Sound levels throughout the cylinder were measured by a moveable microphone traverse. The interior levels were found to be attenuated at most locations between 10 and 35 dB by the active vibration control with a constant control amplitude and phase, i.e., global attenuation was achieved. The physical mechanisms behind the effect are discussed. The new method shows much potential for reduction of propeller interior noise in aircraft, without the penalty of large added weight.

Abstract: The noise transmission characteristics of test panels and acoustic treatments representative of an aircraft sidewall are experimentally investigated in the NASA Langley Research Center transmission loss apparatus. The test panels were built to represent a segment sidewall in the propeller plane of a twin-engine, turboprop light aircraft. It is shown that an advanced treatment, which uses honeycomb for structural stiffening of skin panels, has better noise transmission loss characteristics than a conventional treatment. An alternative treatment, using the concept of limp mass and vibration isolation, provides more transmission loss than the advanced treatment for the same total surface mass. Effects on transmission loss of a variety of acoustic treatment materials (acoustic blankets, septa, damping tape, and trim panels) are presented. Damping tape does not provide additional benefit when the other treatment provides a high level of damping. Window units representative of aircraft installations are shown to have low transmission loss relative to a completely treated sidewall.

Abstract: An ornithopter aircraft has at least a fuselage and four rigid wings which are tandem mounted, in pairs, on opposite sides of the fuselage, in what might be called a "dragonfly configuration". The forward wing in a first of the tandem pairs on one side of the fuselage beats upwardly simultaneously with the diagonally opposed rear wing in the tandem pair on the opposite side of the fuselage, while the remaining two wings are beating downwardly. Then, the wings reverse their direction of travel. The previously upwardly moving wings beat downwardly while the previously downwardly moving wings beat upwardly. The pitch of the wings are varied throughout the beat to produce lift on the downstroke and minimum air resistance on the upstroke, considering the forward speed of the aircraft or the lack thereof. The pitch of the wings are set at the sink angle of a glider wing flying at the same speed. A differential between the picth or stroke of the wings on opposite sides of the fuselage controls direction and banking of the ornithopter.

Abstract: This paper presents the results of an analytical study aimed at predicting the aeromechanical stability of a helicopter in ground resonance, with the inclusions of aerodynamic forces. The theoretical results are found to be in good agreement with the experimental results, available in the literature, indicating that the coupled rotor/fuselage system can be represented by a reasonably simple mathmatical model.

Abstract: The X-wing provides a VTOL aircraft which has a low disc loading hover capability, similar to a conventional helicopter, combined with a high subsonic cruise speed capability. As a result, it hovers with low fuel flow rates which make extended hover duration missions practical. Its low hover power requirements also permit hovering and low speed flight on only one engine in a high speed twin engine aircraft design. The NASA DARPA Sikorsky RSRA X-wing program developed flightworthy X-wing hardware. All design activity and the majority of its component fabrication is completed. A design study was performed on an X-wing concept demonstrator aircraft which is based on the RSRA X-wing components, combined with two MTE engines and a new fuselage.

Abstract: Aerodynamic flow conditions at the air intake of gas turbine engines, espally aircraft engines, comprising boundary layer deflectors are improved by movably arranging the boundary layer deflectors so that they may be retracted flush into the aircraft fuselage during take-off and low speed flight, and may be extended to an operating position during high speed flight. A similar retractable boundary layer deflector may be provided for an air intake for an auxiliary engine, which operates intermittently to drive auxiliary devices of an aircraft. When the auxiliary engine operates the air intake is open and the deflector is extended thereby improving the intake air flow conditions. When the auxiliary engine is shut down, the air intake is covered by a control flap and the deflector is retracted flush into the fuselage. These features provide optimal high- and low-speed flow conditions at the air intakes of the primary propulsion plant and of the auxiliary engine of an aircraft, and undesirable drag is eliminated or at least substantially reduced.

Abstract: Performance features and possible design configurations for oblique wings which would lower the supersonic drag characteristics of fighter aircraft are described. The oblique shape reduces wave drag in proportion to the wing length, and can achieve reductions by factors of 4-16. Variable sweep would be accompanied by fuel pumping to maintain an acceptable center of gravity and tail loading. The lift forces remain balanced and thus remove fuselage loads considerations. Extra tail loading to handle longitudinal moments of center-of-lift shifts with variable sweep aircraft are also eliminated. However, oblique wings couple the roll, pitch and yaw axes and require digital controls to provide balancing forces. NASA has initiated a four phase design, construction and flight test program to examine the flight envelope of oblique wings on the F-8 aircraft.

Abstract: A simplified analytical model is used to study the principal mechanisms at work in propeller noise source radiation, fuselage response, and the behavior of the coupled inner acoustic field, in order to control low frequency sound in aircraft interiors. Both active and passive methods of noise control are comparatively evaluated in light of the transmission mechanisms. Fuselage vibrational response is noted to be dominated by only a few lower order circumferential modes.

Abstract: A method was developed which allows the fully coupled calculation of fuselage and rotor airloads for typical helicopter configurations in forward flight. To do this, an iterative solution is carried out based on a conventional panel representation of the fuselage and a blade element representation of the rotor where fuselage and rotor singularity strengths are determined simultaneously at each step and the rotor wake is allowed to relax (deform) in response to changes in rotor wake loading and fuselage presence. On completion of the iteration, rotor loading and inflow, fuselage singularity strength (and, hence, pressure and velocity distributions) and rotor wake are all consistent. The results of a fully coupled calculation of the flow around representative helicopter configurations are presented. The effect of fuselage components on the rotor flow field and the overall wake structure is detailed and the aerodynamic interference between the different parts of the aircraft is discussed.

Abstract: Theoretical formulations have been developed to describe the transmission of reverberant sound through an infinite, semi-infinite and a finite double panel structure. The model incorporates the fundamental resonance frequencies of each of the panels, the mass-air-mass resonances of the structure, the standing wave resonances in the cavity between the panels and finally the coincidence resonance regions, where the exciting sound pressure wave and flexural waves of each of the panels coincide. It is shown that phase cancellation effects of pressure waves reflected from the cavity boundaries back into the cavity allows the transmission loss of a finite double panel structure to be approximated by a finite double panel mounted in an infinite baffle having no cavity boundaries. Comparison of the theory with high quality transmission loss data yields good agreement in the mass-controlled frequency region. It is shown that the application of acoustic blankets to the double panel structure does not eliminate the mass-air-mass resonances if those occur at low frequencies. It is concluded that this frequency region of low noise transmission loss is a potential interior noise problem area for propeller driven aircraft having a double panel fuselage construction.

Abstract: An arrangement for controlling the air flow past an aerodynamic profile, (such as an aerofoil surface of a wing, tail plane, or fuselage of an aircraft or missile, a propeller, helicopter rotor, or blading of a turbo-engine fan or wind impeller) by providing a number of orifices of small diameter spaced apart in the profile surface in the longitudinal direction of the profile transversely to the direction of air flow over the profile and to the chordwise direction of the profile. Fluid is either discharged or drawn in through the orifices perpendicularly to the air flow in the region of separation of laminar flow. The orifices have a coefficient of flow (i.e., either a discharge or an intake coefficient) CQ =(V/FV) of up to 10-4 for a Reynolds Number less than 3×106, where v=volume flow of the fluid through the discharge orifices, F=profile area and V=velocity of the air stream flowing over the wing or blade. The invention is particularly concerned with controlling the air flow over an aerofoil profile of a wing of a slowly flying aircraft, such as a glider or light aircraft.

Abstract: A nonlinear method based on the full potential equation in conservation form, cast in an arbitrary coordinate system, has been developed to treat predominantly supersonic flows with embedded subsonic regions. This type of flow field occurs frequently near the fuselage/canopy junction area and wing leading-edge regions for a moderately swept fighter configuration. The method uses the theory of characteristics to accurately monitor the type-dependent flowfield. A conservative switching scheme is developed to handle the transition from the supersonic marching algorithm to a subsonic relaxation procedure, and vice versa. An implicit approximate factorization scheme is employed to solve the finite differenced equation. Results are shown for a few configurations, including a wing/body/wake realistic fighter model having embedded subsonic regions.

Abstract: A helicopter includes missile storage means (29) adapted to store a plurality of missiles (26) laterally within a fuselage (28) and missile launcher means (40) located at the end of a lateral outrigger (27) for pivotal movement about a generally vertical axis (42) between a loading position generally laterally of the fuselage and a launch position generally parallel with a longitudinal centreline of the fuselage. In the illustrated embodiment, the launcher means comprises an elongated support beam (45) for selective engagement by a longitudinally extending tubular member (32) attached to each missile and includes transfer means (51, 52, 60, 61) for transferring a missile between the storage means and the launcher means when the latter is in said loading position.

Abstract: A helicopter includes missile supporting means movable from an operational position to a stowed position in which the missiles 17 are located within the confines of the helicopter fuselage In one embodiment the missile supporting means is pivotally attached to the fuselage about a generally horizontal longitudinally extending hinge for movement through an opening in the fuselage between the operational and stowed positions A closure member 16 is provided to automatically close the opening when the missile supporting means is in the closed position and conveniently the missile supporting means may be integral with the closure member

Abstract: An aircraft (1) has a boom (2) pivoted at its forward end to the underside of the aircraft's fuselage (4) and having at its rear end a scoop (11) connected via a duct (10) in the boom (2) to a water storage tank (14) mounted in the fuselage (4). When the aircraft (1) is flying closely above the surface (35) of a body of water (36), means (7,8) for lowering the boom may be used to dip the scoop (11) into the water to thereby refill the tank (14). The water may be jettisoned from the aircraft (1) by opening a dump door (15) built into the tank (14).

Abstract: Various wind tunnel flow visualization techniques have revealed the presence of striations on the surface of both simple geometries and complex winged entry configurations over a wide range of test conditions. The striations are attributed to streamwise vortices embedded in the boundary layer which produce locally enhanced heating levels. Mechanisms for such vortex formation are suggested based on observation of their occurrence on the simple configurations. These mechanisms are then related to striations observed in the wind tunnel on windward surfaces, swept wing leading edges and on the side fuselage of winged entry bodies. Flight data obtained on the side fuselage of ASSET and Shuttle support the conclusions derived from wind tunnel results. Quantitative criteria for the appearance of striations resulting from some of these mechanisms are presented. It is shown that the presence of embedded vorticity may be a consideration in the design of thermal protection systems for advanced winged entry vehicles.