Abstract: The roll-plane radiation patterns of on-aircraft antennas are analyzed using high-frequency solutions. This is a basic study of aircraft-antenna pattern performance in which the aircraft is modeled in its most basic form. The fuselage is assumed to be a perfectly conducting circular cylinder with the antennas mounted near the top or bottom. The wings are simulated by arbitrarily many-sided flat plates and the engines by circular cylinders. The patterns in each case are verified by measured results taken on simple models as well as scale models of actual aircraft.

Abstract: An in-flight aircraft recovery system utilizing an inflatable wing of generally rectangular planform configuration stowed in a normally collapsed condition in a compartment located on the upper portion of the fuselage in the vicinity of the plane''s center of gravity. Upon deployment, the compartment''s cover is ejected and a first parawing type pilot chute lifts a container including the inflatable wing from the compartment and above the tail section of the aircraft. The first pilot chute is jettisoned along with the container after the suspension lines are fully extracted and a second pilot chute of a similar parawing configuration attached to the wing is deployed which positions the inflatable wing above the aircraft with the wing then being inflated by means of a turbine driven compressor mounted on the airfoil surface. The inflatable wing when inflated comprises a rectangular wing including control surfaces in the form of controlled flaps at the wing trailing edge. The wing is connected to the airplane by means of a plurality of suspension lines which are attached to respective rotatable reels. The reels are further controlled for providing selective unreeling and braking of the lines during wing deployment and for subsequently altering not only the angle of attack of the inflated wing, but also the flaps so that the inflated wing flys the aircraft to a predetermined destination either by means of remote pilot control or beacon ground control.

Abstract: The contour adapted passenger loading ramp of the present invention includes, generally, a tunnel defining a walk-way leading to a loading area for receiving a plurality of different models of aircraft having various longitudinal contours. The tunnel terminates at its outer extremity in an access opening having a canopy thereover which is selectively extendable to have its roof engaged with the aircraft fuselage and a contour adaptor is mounted from such fuselage and includes a cover plate pivotally mounted from the canopy and having its free end selectively extendable to accommodate various degrees of fuselage curvature to mate with such fuselage and maintain a continuous ramp roof from the airplane terminal to the entry way to the airplane itself.

Abstract: An aerodynamically balanced high-lift aircraft wherein the problems of large nose-down pitching moments generated by the flap high-lift forces, the loss of trim lift during high-lift flight and the yawing moments caused by the loss of an engine are solved without the use of large horizontal and vertical tails. Also the level of sideline and forward radiated noise is reduced without mechanical apparatus. In the present aircraft, a wing is carried by and bounded on the tips by spaced-parallel fuselages; horizontal tails are mounted only onto the outboard surfaces of the wing-tip fuselages, the centroid-of-lift of the high-lift flaps is located substantially at the center-of-gravity of the aircraft and the exhausts of the engines are emitted in the vertical plane of symmetry of the aircraft. In the aircraft of the present invention, yawing moments occuring during flight with an engine inoperative will be reduced; the horizontal tails will carry an upload and contribute positive trim lift; large nosedown pitching moments generated by the high-lift flaps will be minimized and noise levels will be reduced.

Abstract: A simulated non-compressive type fan jet airplane with a cage at the center of the fuselage and a propeller in the cage that draws air rearwardly to thrust the airplane forward. The propeller is driven by a small electric motor and rechargeable electric batteries that are located in the fuselage of the airplane.

Abstract: A foldable canard assembly for aircraft which have a device for effecting artificial stability and flight control apparatus. The canard assembly comprises a pair of wing units extendable laterally of the aircraft and adjustable about an axis longitudinally arranged of the aircraft to a position extending downwardly therefrom wherein said wing units cooperate together to form a single fin. The axis is preferably located at the bottom of a forward portion of the aircraft fuselage and the wing units are adjustable as desired between fully extended and fully folded positions. A control device is also provided for modifying the horizontal stabilizer surfaces of the tail unit to modify the lift generated thereby in order to compensate for the lift of the canards and thereby maintain longitudinal vertical stability of the aircraft.

Abstract: A definition is given of an idealized fuselage sidewall structure and a simplified analytical model for determining acoustical transmission from the exterior to the interior of a fuselage was constructed. The representation of the sidewall structure chosen for the analytical model excludes complicating effects such as cabin pressurization, acoustic transmission through windows or door seal leaks, aerodynamic excitation, and structural vibration excitation of the fuselage skin.

Abstract: An aircraft includes fuselage flaps which are normally retracted and conform to the configuration of the fuselage and which may be extended outwardly of the fuselage to increase altitude control. The fuselage flaps comprise a first flap positioned rearwardly of the longitudinal center of gravity and hinged on its lower edge about an axis extending upwardly and rearwardly relative to the longitudinal axis of the fuselage. The fuselage flaps also include second and third flaps having their hinge axes aligned with the longitudinal axis of the fuselage and positioned fore and aft of the longitudinal center of gravity of the fuselage. The fuselage flaps further include a fourth and fifth flap aligned with the longitudinal axis of the fuselage at about the longitudinal center of gravity and the fifth flap being hinged to the rear portion of the fourth flap.

Abstract: Recent developments in computer-aided structural design indicate a need for computerized structural analysis techniques which are efficient for the repetitive analysis of large complex structures undergoing design modifications. This paper describes such a technique based on a Taylor series approach. Results are presented for an idealized aircraft fuselage midsection to demonstrate the efficiency and accuracy of the technique. The results show that satisfactory analyses of modified structures may be obtained with the proposed technique, even for large changes in member sizes, for only a small fraction of the computational cost of a full reanalysis.

Abstract: An aircraft having a composite configuration comprising a conventional fuselage forebody portion which symmetrically and variformly blends into a relatively wide, substantially flat beaver-tail-like afterbody, with the afterbody having an integral pitch trimming camber and a composite pitch control device and air brake, hinged transversely thereacross and forming at least a part of its trailing edge The beaver-tail afterbody is geometrically blended into and joined with the after portion of the conventional forebody through the medium of an intervening medial body defined by a geometric transition piece The composite aircraft is further capable of fixedly accommodating any type of conventional wing configuration without the necessity for any adaptative, structural change or modification and also carries conventional propulsion units and control surface components

Abstract: An aircraft of either stable or unstable configuration having all movable wing tip surfaces for rapid increases in lift and subsequent aircraft longitudinal pitch control results from positioning of the wing tip surfaces for positive pitching moments which are then trimmed by trailing edge elevators to produce beneficial positive trim lift forces and having vertical control surfaces to obtain direct side forces for superior control of the aircraft. The aircraft herein has the unique capability of fuselage pointing in the horizontal and vertical axes without change in the flight path as would be the case with conventional aircraft.

Abstract: A radar-absorbing shield for aircraft engines of the type having inlet cavities comprising a geometrically shaped piece of radarabsorbing material which is movably attached to the fuselage of the aircraft near the engine inlet. The shield can be moved under and in front of the inlet to shield the cavity from waves from surface radar, or it can be moved so that it is flush with the fuselage when high maneuverability is required. A shield may be placed near each engine on the aircraft.

Abstract: An automatic temperature control system that maintains passenger comfort levels in aircraft when flying at high altitude. A fuselage skin sensor relays significant temperature changes of the ambient air about the plane to the system temperature controller. Heat transfer into the cabin is thereby smoothly compensated and controlled to predetermined level, during sustained operation in very cold environments. The amount of compensation is adjustable; its operation and control, close with comfortable results.

Abstract: Method and apparatus for attitude control of jet aircraft. In the method, a jet deflection device responds to a variation from the desired attitude in the direction to correct same. Such control is followed by a change in aerodynamic trim control device for effecting correction in the same direction as that effected by the jet deflection and a return of the jet deflection to its initial position. Preferably, the aerodynamic control follows the jet control by a sufficient period of time to minimize its tendency to respond to temporary disturbances such as wind gusts or control inaccuracies. The apparatus aspects of the invention comprise aerodynamic control surfaces which may be canards at the forward end of the fuselage or a sliding wing, together with control and time delay circuitry for adjusting the position of same in response to a variation between actual attitude and desired attitude and preferably including time delay circuitry for effecting the aerodynamic trim correction control a determinable period of time following the jet deflection control.

Abstract: Boundary-layer and local friction data for Mach numbers up to 2.5 and Reynolds numbers up to 3.6 x 10 to the 8th power were obtained in flight at three locations on the XB-70-1 airplane: the lower forward fuselage centerline (nose), the upper rear fuselage centerline, and the upper surface of the right wing. Local skin friction coefficients were derived at each location by using (1) a skin friction force balance, (2) a Preston probe, and (3) an adaptation of Clauser's method which derives skin friction from the rake velocity profile. These three techniques provided consistent results that agreed well with the von Karman-Schoenherr relationship for flow conditions that are quasi-two-dimensional. At the lower angles of attack, the nose-boom and flow-direction vanes are believed to have caused the momentum thickness at the nose to be larger than at the higher angles of attack. The boundary-layer data and local skin friction coefficients are tabulated. The wind-tunnel-model surface-pressure distribution ahead of the three locations and the flight surface-pressure distribution ahead of the wing location are included.

Abstract: Onboard cargo loading means for aircraft. There is provided an upper frame means depending from the aircraft fuselage at a point adjacent a cargo hatch therein. A lower frame structure is movable up and down on the upper frame structure and a lift platform is supported for up and down movement on and along the lower frame structure, with the upper end of the lower frame structure even in its lowermost position overlapping the lower end of the upper frame structure. First and second rollers are placed at the upper and lower end of the lower frame structure. A third roller is placed at the lower end of the upper frame structure and drive sprockets with prime mover means for same are provided elsewhere, preferably on the upper end of the upper frame structure. A chain wraps around said sprocket, under said third roller, over said second roller, under said first roller and back to said sprocket, being fastened at a convenient point therealong to said platform. Thus, energizing of said prime mover first lifts said platform to the upper end of the lower frame structure, then lifts the entirety of said platform and lower frame structure until it fully overlaps the upper frame structure and the platform is flush with the floor of the cargo space of the aircraft. The arrangement of rollers is such that the relative position of the upper and lower frame structures can change with respect to each other without disturbing its operability at any time during a loading operation, particularly if the aircraft sinks lower in response to such loading. In a further development of the invention, means are provided for moving the entire platform supporting structure into the aircraft fuselage for storage or out therefrom for a loading operation. A power operated lever is supported within the aircraft fuselage below the floor of the cargo space and the driven end thereof is connected through one end of a centrally supported lever to an upper end of the platform supporting frame structure. Application of power to said lever will lift said frame structure into the fuselage. The frame structure also carries a brace which in operating or extended position of the loading mechanism bears against aircraft fuselage for steadying said loading mechanism. Linkage is provided which is connected to the other end of said last above-mentioned lever for automatically collapsing said brace against the remainder of said frame structure as the lifting mechanism is retracted into the aircraft fuselage.

Abstract: A cargo ramp system suitable for use with the rear fuselage of low-wing jet transports. The ramp includes a forward ramp section hinged to the aft end of the fixed cargo floor and an aft ramp section pivotally mounted on the fuselage structure. The two sections form a unitary ramp when extended. In flight, the forward section forms an integral part of the cargo floor and the aft section forms an integrap part of the fuselage. During unloading or loading, the forward section is movable to selectively align either with the main cargo floor or with the main ramp section.

Abstract: An aircraft designed to afford passengers an unrestricted view for "sightseeing" or other purposes. Passenger accommodation extends to the foremost part of a fuselage and is enclosed by a cabin transparency extending to cover substantially the entire field of view of passengers from a fixed position. The pilot of the aircraft is accommodated above and behind the passengers, the aircraft's powerplant also being disposed behind the passenger accommodation.

Abstract: A sound suppression apparatus for a jet engine mounted within the tail portion of an aircraft fuselage wherein a pair of aerodynamic surfaces extend outwardly and upwardly on each side of the fuselage and are mounted for movement about a lateral axis for varying their angle of incidence. The aerodynamic surfaces or panel members from a rear perspective view, form a "V" or "U" shaped configuration, with their point of intersection with the fuselage below the main stream of engine exhaUst flow when a variable exhaust nozzle is in its takeoff and landing position or sound suppression mode. By varying the angle of incidence of the aerodynamic surfaces, the function to direct the streamwise airflow, adjacent to the external surface of the fuselage, sideways into the exhaust nozzle flow as to to come in underneath it, as an intermediate fluid stream. This intermediate airflow, causes a mixed boundary layer region beneath the engine exhaust flow which suppresses the downward radiation of the engine exhust noise. Normally, the noise in the exhaust flow would follow the surface contour of the exhaust duct or any extension thereof and spill over the end resulting in a substantial downward angle or come of noise radiation. However, through the introduction of the intermediate flow, a mixed boundary layer flow region is formed throughout the length of the exhaust duct and any undersurface extension thereof, which boundary layer detaches the noise flow from following the surface and provides for greater effectivity of noise deflecting surfaces or shields positioned about the exhaust flow.

Abstract: The upper-surface blown (USB) flap as a powered-lift concept has evolved because of the potential acoustic shielding provided when turbofan engines are installed on a wing upper surface. The results from a wind tunnel investigation of a large-scale USB model powered by two JT15D-1 turbofan engines are-presented. The effects of coanda flap extent and deflection, forward speed, and exhaust nozzle configuration were investigated. To determine the wing shielding the acoustics of a single engine nacelle removed from the model were also measured. Effective shielding occurred in the aft underwing quadrant. In the forward quadrant the shielding of the high frequency noise was counteracted by an increase in the lower frequency wing-exhaust interaction noise. The fuselage provided shielding of the opposite engine noise such that the difference between single and double engine operation was 1.5 PNdB under the wing. The effects of coanda flap deflection and extent, angle of attack, and forward speed were small. Forward speed reduced the perceived noise level (PNL) by reducing the wing-exhaust interaction noise.

Abstract: The unsteady aerodynamics of the 040A orbiter have been explored experimentally. The results substantiate earlier predictions of the unsteady flow boundaries for a 60 deg swept delta wing at zero yaw and with no controls deflected. The test revealed a previously unknown region of discontinuous yaw characteristics at transonic speeds. Oilflow results indicate that this is the result of a coupling between wing and fuselage flows via the separated region forward of the deflected elevon. In fact, the large leeward elevon deflections are shown to produce a multitude of nonlinear stability effects which sometimes involve hysteresis. Predictions of the unsteady flow boundaries are made for the current orbiter. They should carry a good degree of confidence due to the present substantiation of previous predictions for the 040A. It is proposed that the present experiments be extended to the current configuration to define control-induced effects. Every effort should be made to account for Reynolds number, roughness, and possible hot-wall effects on any future experiments.

Abstract: A toy airplane featuring a control member rotatably mounted to the fuselage that allows the airplane to be flown in both the clockwise and counterclockwise directions, either right side up or upside down. The control member includes a swivel mounting that it located within an open channel on the underside of the fuselage and between the wings of the airplane.

Abstract: A mathematical model for a real time simulation of a tilt rotor aircraft was developed. The mathematical model is used for evaluating aircraft performance and handling qualities. The model is based on an eleven degree of freedom total force representation. The rotor is treated as a point source of forces and moments with appropriate response time lags and actuator dynamics. The aerodynamics of the wing, tail, rotors, landing gear, and fuselage are included.

Abstract: : This report identifies fuel temperature levels and contributors to fuel temperature rise or decrease at each step of fuel handling or usage from ground bulk storage to engine combustor. The program consisted of a literature search and review of available information on fuel temperatures in bulk storage and on-loaded aircraft tanks, flight profile effects on fuel tank fuel temperature, and fuel temperature changes resulting from aircraft and engine heat loads in flight. The study program was supplemented by MINEX thermal stability tests on JP fuels by the USAF. The results indicate a low incidence of bulk storage or refueling fuel temperatures above 95 F. Aircraft wing tank fuel temperatures, in static ground soak and in flight, follow closely the changes in local ambient or free stream total temperature. Fuselage or body tank fuel temperatures in static ground soak or in flight, have a gradual change with respect to large differences in local ambient or free stream temperature. The sources and levels of heat loads from aircraft and engine are established. Fuel temperatures are significantly influenced by the power requirements, the environment, and the design for thermal integration of the engine and aircraft fuel systems. The results of the MINEX thermal stability tests show a wide range in relative quality level for the fuels tested. The overall results indicate that aircraft and engine systems can be designed to operate in the Mach 3 range using present primary type fuels and state-of-the-art fluid system components.

Abstract: Antennas in the proximity of complicated conducting structures such as aircraft, submarines, missiles, and satellites have, in general, analytically unpredictable impedance characteristics due to coupling effects between the antennas and these structures. As a model of the fuselage of these conducting structures, a conducting cylinder of finite length is considered and an analytic-numerical technique is used to compute the impedance characteristics of a nearby linear dipole of arbitrary length.

Abstract: An experimental program was performed in which the individual performance of multiple VTOL model lift fans was measured. The model tested consisted of three 5.5 in. diameter tip-turbine driven model VTOL lift fans mounted chordwise in a two-dimensional wing to simulate a pod-type array. The performance data provided significant insight into possible thrust variations and losses caused by the presence of cover doors, adjacent fuselage panels, and adjacent fans. The effect of a partial loss of drive air supply (simulated gas generator failure) on fan performance was also investigated. The results of the tests demonstrated that lift fan installation variables and hardware can have a significant effect on the thrust of the individual fans.

Abstract: An experimental investigation was conducted at Mach 6 to determine the hypersonic aerodynamic characteristics of an all-body, delta-planform, hypersonic research aircraft (HYFAC configuration). The aerodynamic characteristics were obtained at Reynolds numbers based on model length of 2.84 million and 10.5 million and over an angle-of-attack range from minus 4 deg to 20 deg. The experimental results show that the HYFAC configuration is longitudinally stable and can be trimmed over the range of test conditions. The configuration had a small degree of directional stability over the angle-of-attack range and positive effective dihedral at angles of attack greater than 2 deg. Addition of canards caused a decrease in longitudinal stability and an increase in directional stability. Oil-flow studies revealed extensive areas of separated and vortex flow on the fuselage lee surface. A limited comparison of wind-tunnel data with several hypersonic approximations indicated that, except for the directional stability, the tangent-cone method gave adequate agreement at control settings between 5 deg and minus 5 deg and positive lift coefficient. A limited comparison indicated that the HYFAC configuration had greater longitudinal stability than an elliptical-cross-section configuration, but a lower maximum lift-drag ratio.

Abstract: : The report presents the results of a test program to measure the near-field wake behind the wing of a test aircraft, out of the ground effect, and to obtain vortex trajectory data on trailing vortices generated within the ground effect. The test aircraft was a high-lift L-19 aircraft incorporating a distributed-suction boundary layer control system. The wake measurement probe was supported from the fuselage of the test aircraft and was mounted on a trolley which moved along a boom structure. The velocity measurements were made with a six-element hot-film anemometer. The near field wake data were obtained in two measurement planes located 0.64 and 5.10 feet behind the wing. Data were obtained at zero and fullflap deflection and true speeds of 46, 66, and 85 mph. The data are presented as contours of constant normalized vorticity, nondimensional vertical velocity, and non-dimensional longitudinal velocity increment within the measurement planes. (Modified author abstract)