Abstract: A cruciform wing structure for a solar powered aircraft is disclosed. Solar cells 28 are mounted on horizontal wing surfaces 20, 22. Wing surfaces 24, 26 with spanwise axis perpendicular to surfaces 20, 22 maintain these surfaces normal to the sun's rays by allowing aircraft to be flown in a controlled pattern at a large bank angle. The solar airplane may be of conventional design with respect to fuselage, propeller and tail, or may be constructed around a core 70 and driven by propeller mechanisms 75, 76, 77, and 78 attached near the tips of the airfoils.

Abstract: The invention relates to aircraft wings, wherein the fairing element between the leading edge of a wing and the fuselage comprises a streamlining flap movable between a folded position in which it is housed in a recess in the fairing element and forms a part of the leading edge and the part adjacent the lower surface of said fairing element, and an extended position in which it ensures continuity of the leading edge between the fuselage and the high-lift spoiler likewise in extended position. Moreover, there is a communication between the housing of the streamlining flap and the housing of the high-lift spoiler so that, when these two flap elements are in their extended positions, the air penetrating in the housing of the streamlining flap disposed on the lower surface side emerges through the housing of the high-lift spoiler disposed on the upper surface side, with the result that there is an increased upper surface blowing.

Abstract: One of the dominant source-path combinations for cabin noise in light twin-engine aircraft is propeller noise being transmitted through the fuselage sidewall. This source-path was investigated and candidate sidewall add- on treatments were installed and tested using both an external sound source and the propeller in ground static engine runs. Results indicate that adding either mass or stiffness to the fuselage skin would improve sidewall attenuation and that the honeycomb stiffness treatment provided more improvement at most frequencies than an equal amount of added mass. It is proposed that double-wall construction in conjunction with skin stiffening should provide a good weight-efficient combination for the aircraft studied. NE of the principal source-path combinations of cabin noise in light, twin-engine aircraft is propeller noise transmitted through the fuselage sidewall. Improved methods of controlling this cabin noise are needed to provide a comfortable passenger environment, while at the same time controlling aircraft weight and fuel consumption. Lighter weight noise control methods are needed to replace traditional approaches which have relied largely on relatively heavy damping and mass treatments. A number of approaches have been investigated for reducing cabin noise for this type of aircraft. Flight tests indicated interior noise can be reduced about 3.5 dB(A) by a reduction of engine rpm in an aircraft with variable pitch propellers.1 Design of propeller configurations is being in- vestigated as a means of reducing the noise generated at the source.2 Theoretical prediction methods for sidewall noise transmission have been developed to aid the search for noise- resistant sidewall structures. Theoretical analysis of interior noise transmission has included mechanical analogy models, rigid-stiffener/flexible-panel models, and more complex flexible-stiffener/flexible-panel models. 3"5 The analyses have been compared with laboratory test data for verification and have been used to examine a number of candidate noise control treatments including variations of skin thickness, stiffener stiffness, and structural damping, and addition of damping, mass, and honeycomb panel stiffening. Previous work has not included evaluation of candidate noise control treatments in an experimental situation using an actual aircraft. Such studies are needed to evaluate and compare candidate treatments, and to guide further development of noise control treatments. The purpose of this paper is to describe an experimental program of evaluation of three noise control treatments. The work is focused on added stiffness in the form of honeycomb panels. Also, two mass treatments are included for comparison. The tests were carried out using a light twin-engine aircraft (Fig. 1). Can- didate treatments were developed using the aircraft with a horn noise source in the laboratory. The performance of the stiffness treatment was verified using ground static runs of the aircraft engines. The laboratory portion of this investigation is described in Ref. 6.

Abstract: The feasibility of the propfan relative to the turbofan is summarized, using the Douglas DC-9 Super 80 (DS-8000) as the actual operational base aircraft. The 155 passenger economy class aircraft (31,775 lb 14,413 kg payload), cruise Mach at 0.80 at 31,000 ft (8,450 m) initial altitude, and an operational capability in 1985 was considered. Three propfan arrangements, wing mounted, conventional horizontal tail aft mounted, and aft fuselage pylon mounted are selected for comparison with the DC-9 Super 80 P&WA JT8D-209 turbofan powered aircraft. The configuration feasibility, aerodynamics, propulsion, structural loads, structural dynamics, sonic fatigue, acoustics, weight maintainability, performance, rough order of magnitude economics, and airline coordination are examined. The effects of alternate cruise Mach number, mission stage lengths, and propfan design characteristics are considered. Recommendations for further study, ground testing, and flight testing are included.

Abstract: T wing-fuselage problem is investigated by means of the Navier-Stokes equations incorporated with an approximate turbulence model. The numerical solution yields a reasonable global agreement with experimental data in static and impact pressure distributions. However, in order to better describe the flowfield structure near the leading edge of the wing, an alternative choice of coordinate system is required.

Abstract: The structural, manufacturing, and service and environmental considerations that could impact the design of composite fuselage structure for commercial transport aircraft application were explored. The severity of these considerations was assessed and the principal design drivers delineated. Technical issues and potential problem areas which must be resolved before sufficient confidence is established to commit to composite materials were defined. The key issues considered are: definition of composite fuselage design specifications, damage tolerance, and crashworthiness.

Abstract: Combined analytic and experimental activities were performed to evaluate an advanced structural design concept for advanced composite fuselage panels suitable for V/STOL aircraft. An existing sonic fatigue analysis procedure was then evaluated. Both flat and slightly curved multibay cross-stiffened panels with graphite-epoxy skins were designed, analyzed, fabricated, and tested. The panels simulated skin-frame-stringer aircraft structure. The joints of the substructure and skin were achieved in a single-cure autoclave operation without an adhesive. The substructure consisted of graphite-epoxy hat-section stringers and glass-epoxy J-section frames. Some panels were tested in a combined acoustic-thermal environment, and others were tested in an acoustic environment at room temperature. The sonic fatigue tests were performed in a progressive wave acoustic test chamber with broadband excitation at approximately 163.5-dB overall sound pressure level. The sonic fatigue lives of the test panels exceeded the predicted lives obtained with existing semiempirical methods developed for graphite-epoxy and aluminum alloy panels featuring other design concepts and manufacturing processes.

Abstract: An aircraft having a fuselage provided with an internal duct extending longitudinally therethrough to provide an internal wing for the craft, the internal duct having the forward end open for receiving an air stream therethrough and the aft end thereof open for discharge of the air stream therefrom, the internal contour of the duct being alterable in accordance with required operational conditions for the flight of the craft, and a plurality of control flaps and/or vanes provided at the aft end of the duct for proving operational controls for the craft in the manner of a more conventional external wing craft.

Abstract: An aircraft having a fuselage and a pair of forward-swept wings includes a wing carry through extending transversely through the fuselage joining the wings together to form a unitary wing structure, two hinge assemblies, each located adjacent a different one of the wings for rotatably mounting the wing structure to the fuselage and located forwardly of the wing carry through and on a spanwise axis intersecting the aerodynamic centers of the wings, and two actuators, each located rearwardly of the hinge assemblies and extending between the fuselage and a different one of the wings for selectively rotating the wing structure about the hinge assemblies The hinge assemblies attach the wing structure to the fuselage such that the axis of rotation of the wing structure is substantially collinear with the axis intersecting the aerodynamic centers of the wings so that relatively little force is required of the actuators to vary the angle of incidence of the wing structure

Abstract: The NASA Langley Research Center has initiated a broad general aviation stall/spin research program. A rotary balance system was developed to support this effort. This system, located in the Langley spin tunnel, makes it possible to identify airplane aerodynamic characteristics in a rotational flow environment, and thereby permits prediction of spins. This paper presents a brief description of the experimental setup, testing technique, five model programs conducted to date, and an overview of the rotary balance results and their correlation with spin tunnel free-spinning model results. It is shown, for example, that there is a pronounced nonlinear dependency of the aerodynamic moments on rotational rate and that these moments are very configuration dependent. Fuselage shape, horizontal tail, and, in some instances, wing location are shown to appreciably influence the yawing moment characteristics above an angle of attack of 45 deg.

Abstract: A mounting for a helicopter pylon having a plurality of resilient vertical support links (24-27) between the pylon (12) and the helicopter fuselage (13). The links have a composite spring rate for a given pylon weight which establishes a resilient support for the fuselage from the pylon at the thrust required to support said fuselage in normal flight and in which oscillatory vertical force transmitted to fuselage is minimal for acceptable static deflections of the pylon. Physical stops (24e, 24n, 24p) are operable to limit deflection between the pylon (12) and the fuselage (13) at predetermined rotor thrust above and below said normal thrust. Links (24-27) are mounted at the ends of a pair of rigid beams (20, 21 and 22, 23) which are secured at the centers thereof to opposite sides of transmission portion of the pylon (12). A vertical rigid mounting pin (24b) extends downward at the end of each beam. A hollow vertical receptacle (24a) is secured to the fuselage and is coaxial with each pin (24b) and has a reentrant bore (24h) providing a down facing shoulder (24p). A resilient body (24g) couples each pin to its receptacle. An enlarged foot (24e) on each pin operates between the fuselage (13) and shoulder (24p) to limit deflection of pylon (12) and the resilient coupling (24g).

Abstract: This invention is an improvement of my system for supersonic aircraft which provides the required angular momentum reaction to the continuous generation of new lift circulation by vorticity in lieu of dissipative shock waves. This new system relocates the propulsive jet forward utilizing its excess energy to generate this vorticity in a planar sheet below the wing acting as a pressure shield, which continues downstream beyond the wing trailing edge as a vortex flap simulating an extended wing chord. The present invention provides mechanical vortex generating means to augment this free surface mechanism in cases where the jet velocity is insufficient. This invention also incorporates a wide-body fuselage above this pressure shield, and provides further system improvement in propulsion/wing integration.

Abstract: Transonic modified small-disturbance theory has been employed to numerically model the flowfield around wing/fuselage/pylon/store configurations. A fine grid region enclosing the wing/pylon/store is embedded within a global crude grid and a successive crude-fine relaxation is performed. Using an image point concept, the store and the pylon are introduced into an existing wing/fuselage program, thus avoiding excessive additional computer memory requirements. Comparison of results with experiments on the F-5 wing with a pylon/store arrangement is presented showing good agreement. A study of the roles of pylon height, store diameter, pylon span mount location, angle of attack, and Mach number relative to the achievement of optimum LID from beneficial nonlinear interference is presented. In addition, a simplified analytical approach to compute the loading on the store using an "immersion theory" is indicated and validated against experiments.

Abstract: A dropable main fuel tank that comprises a portion of an aircrafts lower fuselage or part of the wing structure contains all nonessential fuel for takeoff or landing. Airfoils, parachute, or rocket motor built into the main fuel tank will separate the main fuel tank from the aircraft quickly when dropped or released in a possible crash situation, thus preventing the aircraft from being consumed by its own nonessential fuel in a crash. A nondroppable auxiliary fuel tank on the aircraft will sustain the aircraft after the main fuel tank has been dropped. The auxiliary fuel tank contains only enough fuel for takeoff and landing. Shackels and parallel tracks hold the droppable main fuel tank in place on the aircraft until the main fuel tank is released for subsequent ejection rearward. In refueling the aircraft an external control panel on the aircraft will release the main fuel tank onto a refueling vehicle for refueling at a nearby fuel dump.

Abstract: An aircraft containing aerodynamic and gyroscopic stability that has a high angle of take-off and landing capabilities, with high speed horizontal powered flight, but is also able to sustain power-off flight with auto-rotation of its multiple extending airfoils. The fuselage has the shape of an inverted saucer with aerodynamic configuration and has an open circular track at its periphery; and riding in this track is a rotary frame and extended airfoil assembly that isin-line with the fuselage. Each extending airfoil contains solid weighted bodies at their tips and this rotary frame and extended airfoil assembly is rotated on the track by an internal power unit. The extending airfoils taper toward their tips and these tips have a knife sharp edge for penetrating the air resistance. Special flaps on the airfoils function for creating additional lift in the downwind quadrant, but in the power-off flight mode, when the rotary frame and extending airfoil assembly is disengaged; the dual purpose flap and trap flap combine to function to trap the slip-stream causing the rotary frame and extended airfoil assembly to rotate. Mounted to the surface of the fuselage are forward thrust engines with rearward extending booms on which are located the tail assembly with flight control surfaces.

Abstract: A tubular member holds a paper, tubular fuselage above a paper wing supported by a plate and below a folding die from which a fuselage presser depends; whereby the fuselage may be attached to the wing and the edges of the wing may be folded into an aerodynamic configuration when the die and tubular member are brought into working association with a pair of upstanding plates which are spaced apart in a manner such that the fuselage and its associated parts may be positioned between said plates while said die folds the wing edges over the upper edges of the plates.

Abstract: : The Army YAH-64 and UH-60A helicopters were studied to determine optimum wind tunnel test procedures and to expand the experimental data base on drag reduction for rotor hubs and pylons. Full- and reduced-scale models of these helicopters were fabricated. The Hub Pylon Evaluation Rig (HPER) and the Generalized Rotor Modeling System (GRMS) were used for the experimental testing, conducted in the NASA/Langley V/STOL wind tunnel. Only the YAH-64 models underwent wind tunnel testing during this contracted effort. Plans are under way for testing of the UH-60A models. All configurations were subjected to viscous analysis using Program DRAG, a configuration modeling program. This effort included evaluation of hub fairings, pylon fences, rotor wake flow, hub rotation, engine air flow, fuselage parasite drag, empennage flow, and stabilators. The DRAG program was validated by correlation of predicted and experimentally obtained surface pressures.

Abstract: Exterior noise was measured on the fuselage of a twin-engine, light aircraft at four values of engine rpm in ground static tests and at forward speeds up to 36 m/s in taxi tests. Propeller noise levels, spectra, and correlations were determined using a horizontal array of seven flush-mounted microphones and a vertical array of four flush-mounted microphones in the propeller plane. The measured levels and spectra are compared with predictions based on empirical and analytical methods for static and taxi conditions. Trace velocities obtained from point-to-point correlations are used to describe the propagating and rotating characteristics of the propeller noise field on the fuselage.

Abstract: : Force determination is a method of obtaining dynamic loads acting on a vehicle in flight. These loads were determined from measured fuselage responses obtained in flight and calibration matrices obtained in a shake test. These forces obtained were verified by ground flying in a hangar and duplicated the responses obtained in flight.

Abstract: A study is performed of rotary balance data, spin tunnel model, radio-controlled (R/C) model, and full-scale flight test results relating to the spinning of light aircraft. A method is presented for predicting steady spin modes using rotary balance data. Differences in spin characteristics of various wing, tail, and fuselage modifications are discussed as well as scale effects. It is concluded that an equilibrium flat spin is governed primarily by the yawing moment coefficient.

Abstract: Aircraft lowering and raising apparatus for loading and unloading purposes wherein the apparatus comprises four supporting structures, two located at a spaced distance at the left and two similarly at the right side of the fuselage. Each structure is horizontally carried within the aircraft and can be extended in a vertical position outside of the aircraft. Each structure includes a hydraulic telescoping piston cylinder arrangement with a ground feeler for extending a tripod carriage which is pivotally connected to the cylinder. Upon operation, the structure moves from its aircraft storage unit outside in vertical position and engages the ground by extension of its telescoping member to support the aircraft, thereupon the landing gears are contracted into the aircraft body, and thereafter the telescoping piston cylinder is contracted so that the aircraft is lowered for cargo handling. Raising of the airplane on its landing gears for takeoff is obtained by reversing the described operations of the supporting structures.

Abstract: The status of ejector development in terms of application to V/STOL aircraft is reported in three categories: aircraft systems and ejector concepts; ejector performance including prediction techniques and experimental data base available; and, integration of the ejector with complete aircraft configurations. Available prediction techniques are reviewed and performance of three ejector designs with vertical lift capability is summarized. Applications of the 'fuselage' and 'short diffuser' ejectors to fighter aircraft are related to current and planned research programs. Recommendations are listed for effort needed to evaluate installed performance.

Abstract: An extensive program has been undertaken to investigate the dynamic behaviour of fuselage structures subject to various impact conditions. Extensive testing of scale model stiffened aluminum sections has been completed for a wide range of wing loads, angles of incidence and impact velocities. Both vertical drop tests and "free-flight" impacts using a pendulum gantry have been studied. Test data have been obtained in terms of maximum structural s trains, g-loads and high speed photographs of the dynamic collapse modes. Based on a finite element model, these cases have also been analysed including elastic/plastic deformation and panel buckling. Some comparative results are then presented together with computer graphics.

Abstract: Full-scale wind-tunnel tests, conducted to determine the effects of three different airfoil sections on the aft-fuselage drag of a low-wing aircraft, are described. The measurements indicate a maximum difference in aft fuselage drag between the three airfoils of about 0.002. Measured changes in the locations of the fuselage pressure contours with airfoil section correlated well with the changes predicted by a three dimensional paneling code. A criterion, based on the inviscid computer code, is then proposed as an indicator for possible adverse viscous interactions at the wing-fuselage juncture.

Abstract: Recent NASA and NASA sponsored research on the prediction and control of propeller and rotor source noise, on the analysis and design of fuselage sidewall noise control treatments, and on the measurement and quantification of the response of passengers to aircraft noise is described. Source noise predictions are compared with measurements for conventional low speed propellers, for new high speed propellers (propfans), and for a helicopter. Results from a light aircraft demonstration program are considered which indicates that about 5 dB reduction of flyover noise can be obtained without significant performance penalty. Sidewall design studies are examined for interior noise control in light general aviation aircraft and in large transports using propfan propulsion. The weight of the added acoustic treatment is estimated and tradeoffs between weight and noise reduction are discussed. A laboratory study of passenger response to combined broadband and tonal propeller-like noise is described. Subject discomfort ratings of combined tone broadband noises are compared with ratings of broadband (boundary layer) noise alone and the relative importance of the propeller tones is examined.

Abstract: A 1/20-size, low-speed flutter model of the SCAT-15F complete airplane was tested on cables to simulate a near free-flying condition. Only the model wing and fuselage were flexible. Flutter boundaries were measured for a nominal configuration and a configuration with wing fins removed at Mach numbers M from 0.76 to 1.2. For both configurations, the transonic dip in the wing flutter dynamic pressure q boundary was relatively small and the minimum flutter q occurred near M = 0.92. Removing the wing fins increased the flutter q about 14 percent and changed the flutter mode from symmetric to antisymmetric. Vibration and flutter analyses were made using a finite-element structural representation and subsonic kernel-function aerodynamics. For the nominal configuration, the analysis (using calculated modal data) predicted the experimental flutter q levels within 10 percent but did not predict the correct flutter mode at the higher M. For the configuration without wing fins, the analysis predicted 16 to 36 percent unconservative (higher than experimental) flutter q levels and showed extreme sensitivity to mass representation details that affected wing tip mode shapes. For high subsonic M, empennage aerodynamics had a significant effect on the predicted flutter boundaries of several symmetric modes.

Abstract: This paper is a summary of an analysis of the ground-effect characteristics of a large-scale twin-engine, tilt-nacelle V/STOL model. The analysis considers data from the flow field beneath the full-scale model, as well as small-scale model test data, and makes comparisons with jet-ground interactions predicted by a computer code. The data from the large-scale test comprise ground-plane surface temperatures, static pressure distribution and wall-jet total-pressure profiles, fuselage undersurface static pressures, and model forces and moments. The results indicate that the near-field flow is more complex than is indicated by either the small-scale uniform jet studies or the computer predictions. The far-field flow characteristics do show some similarity for these three cases.

Abstract: It is noted that most modern fighter aircraft rely on vortex interaction to provide lift enhancement at maneuvering angles of attack. It is shown that the close-coupled horizontal canard in a three-surface configuration provides a control surface which in addition to its other control functions, can be used to optimize this vortex interaction. Attention is given to a study intended to provide a detailed understanding of the aerodynamics of this type of configuration. The discussion examines the results of this investigation and hypotheses are presented to explain the linear and nonlinear aerodynamic phenomena observed.