Abstract: A vertical and short take-off and landing aircraft comprising a fuselage, a set canard wings, a set of lift fan wings, air deflectors, lift wings, and a pusher propeller. The canard wings are attached forward of the center of gravity to the fuselage. The lift fan wings are attached about the fuselage generally about the center of gravity of the aircraft. The lift fan wings comprise a generally circular duct extending vertically through the wing, a multi-bladed fan mounted for free rotation axially in the duct, and a prime mover connected to the fan for selectively applying rotational torque to the fan. The air deflectors are arranged about the lift fan wing in a louver-type of system for directing even flow of air to the fan. The lift wings are attached to the fuselage aft of the center of gravity of the aircraft and generally at a location vertically higher than the lift fan wings. The pusher propeller is connected to the prime mover and attached to the fuselage aft of the lift Jan wings.

Abstract: A floor construction for the upper compartment of an aircraft having an about circular fuselage cross-section. A one piece floor plate provides bending supports and extends over nearly the full inner diameter of the fuselage. A connection is provided between the edge of the floor plate and the fuselage outer structure for transmitting forces directed predominantly in the longitudinal direction of the aircraft. Vertically running planar structural members provide side support elements for the floor plate. The construction results in a decrease of the construction weight as well as of the costs, in particular if the floor plates are provided from fiber reinforced plastic. The connection between the floor plates and the fuselage frame can be provided by rods running at an angle forwardly, which transmit preferably forces acting in the longitudinal direction of the aircraft.

Abstract: Future high-performance aircraft will have high thrust/weight ratios, will be equipped with controlconfigured vehicle technology, will be extremely lightweight, and possibly have the capability of flying at very high angles of attack in the "poststall" region. This paper investigates, by using numerical optimization techniques, whether the poststall capability improves performance for several tactical maneuvers. Specifically, minimum-time turning maneuvers for a variety of boundary conditions and flight-path constraints are computed 1) for aircraft A which has poststall capability and 2) for aircraft B which does not, but is otherwise identical to A. It is concluded that for two combinations of boundary conditions/path constraints, flight time can be reduced if high angles of attack are utilized. In the majority of cases, however, minimum-time maneuvers are flown, load constraints permitting, at or near the maximum lift coefficient.

Abstract: A V/STOL aircraft comprising a fuselage, having three sets of wings that are offset lengthwise and vertically. The center set of relatively small relatively thick wings 34 between the canard 14 and the rear relatively thin relatively large aerodynamic lift wings 18 are equipped with identical lift fans 36 enclosed in the wing by upper and lower movable slotted deflectors 40. The attitude of the slatted deflectors may be varied to transition the present invention from hovering to forward flight, and vice versa. The present invention's lift fans are interconnected by a balanced power distribution system, to insure constant, efficient use of total power and provide symmetrical lift about the aircraft's center of gravity resistibility. In the event of an engine failure, the remaining lift is still properly distributed to maintain symmetrical lift, so as to maintain balance and operational control of the aircraft in the lift mode.

Abstract: A structural weight comparison was made between a new concept wing design, called a "Joined Wing," and a reference conventional wing-plus-horizontal tail (Boeing 727). The joined-wing analysis includes two cases that differ only in minimum gage skin thickness. The comparison was accomplished by constructing finite-element computer models of each wing configuration, analyzing each for optimum skin thickness, then determining the structural weight of each wing. The optimizations were based on a fully stressed design concept using a Von Mises criterion for maximum allowable stress. The joined wing was found to be lighter by 12-22%.

Abstract: A design system for aircraft that allows wings of various different planforms to be mounted interchangeably on a common fuselage. Each of the wings has wing semi-spans joined by a center section torque box that mounts the wing in a cut-out in the aircraft fuselage. The torque box has a quadrilateral structure including front and rear main center section spars. These center section spars have their ends connected to the front and rear main spars of the semi-spans to transmit wing loads into the fuselage. The length of the main center section spars are designed in accordance with the invention such that they can connect with the main spars of the wing semi-spans to position the quarter chord of the wing properly with respect to the center of gravity of the aircraft irrespective of the wing planform. Thus, aft-swept, forward-swept, or straight wings can be routinely interchanged on a common fuselage without effecting the positive static longitudinal stability of the aircraft. Interchangeable wings can be used with a common fuselage to produce either a high or a low wing configuration.

Abstract: A moveable wing aircraft including a quick release, attachment mechanism carrying the aircraft on a bomb rack or other carrier and a mechanism for deploying the moveable wing from its captive carry position to its extended free flight position are disclosed. The aircraft includes an elongate fuselage, a portion of the top surface of which is substantially flat in order to accommodate the moveable wing. The moveable wing is positional between a captive carry position in which it is aligned with the longitudinal axis of the fuselage and an extended free flight position. The single, moveable wing is pivoted around a central point from its captive carry position to its extended free flight position such that it is substantially perpendicular to the aircraft fuselage. The quick release mechanism extends through apertures in the wing in its captive carry position and is spring biased to retract through the wing and into the aircraft fuselage when released from the bomb rack or other carrier. The deployment mechanism includes a spring loaded cable and pulley arrangement and serves to connect the moveable wing to the fuselage and to bias it from its captive carry position to its extended free flight position when activated upon release of the quick release mechanism.

Abstract: An empennage assembly for supersonic aircraft includes longitudinally and rearwardly extending booms (12) mounted upon the wings (16) or fuselage of the aircraft. The booms (12) include aft rotatable sections (22) upon which are mounted larger (18) and smaller (20) tail surfaces. The boom sections (22) are angularly rotated through angular displacements θ such that the dispositions of the tail surfaces (18, 20) are interchanged between their dispositions during low subsonic and high supersonic flight conditions. In this manner, the directional and longitudinal aerodynamic static stability components of the aircraft are rendered substantially constant at an optimum low or near-neutral level of stability in order to enhance the flight maneuverability capabilities of the aircraft throughout the subsonic, transonic, and high supersonic speed ranges. The empennage (10) is also advantageously employed for enhancing the directional stability characteristics of all aircraft under varied angle of attack conditions, and still further, can likewise enhance the lift characteristics of STOL aircraft employing vectored thrust. In connection with the use of the present invention empennage system upon supersonic aircraft, the rotational orientation of the empennage system (10) and its associated boom sections (22) is automatically programmed by suitable feed-back control means (23) in response to sensed changes in Mach number. In a similar manner, the rotational orientation of the empennage system (10) and its associated boom sections (22) would also be automatically programmed by the feed-back control means (23) as a function of angle of attack.

Abstract: Crashworthy floor concepts applicable to general aviation aircraft metal airframe structures were investigated. Initially several energy absorbing lower fuselage structure concepts were evaluated. Full scale floor sections representative of a twin engine, general aviation airplane lower fuselage structure were designed and fabricated. The floors featured an upper high strength platform with an energy absorbing, crushable structure underneath. Eighteen floors were fabricated that incorporated five different crushable subfloor concepts. The floors were then evaluated through static and dynamic testing. Computer programs NASTRAN and KRASH were used for the static and dynamic analysis of the floor section designs. Two twin engine airplane fuselages were modified to incorporate the most promising crashworthy floor sections for test evaluation.

Abstract: A state-variable feedback approach is utilized for active control of rotorcraft vibration. Fuselage accelerations are passed through undamped second-order filters with resonant frequencies at N/rev. The resulting outputs contain predominantly the N/rev vibration components, phase shifted by 180 deg, and are used to drive the blade pitch to cancel this component of fuselage vibration. The linear-quadratic-gaussian (LQG) method is used to design a feedback control system utilizing these filtered accelerations. The design is based on a nine-degree-of-freedom linear model of the Rotor System Research Aircraft (RSRA) in hover and is evaluated on a nonlinear blade-element simulation of the RSRA for this flight condition. The system is shown to essentially eliminate vibrations at N/rev in all axes. The required blade-pitch amplitude is within the capability of conventional actuators at the N/rev frequency.

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: Four six-place, low-wing, twin-engine, general aviation airplane test specimens were crash tested under controlled free flight conditions. All airplanes were impacted on a concrete test surface at a nomial flight path velocity of 27 m/sec. Two tests were conducted at a -15 deg flight path angle (0 deg pitch angle and 15 deg pitch angle), and two were conducted at a -30 deg flight path angle (-30 deg pitch angle). The average acceleration time histories (crash pulses) in the cabin area for each principal direction were calculated for each crash test. In addition, the peak floor accelerations were calculated for each test as a function of aircraft fuselage longitudinal station number. Anthropomorphic dummy accelerations were analyzed using the dynamic response index and severity index (SI) models. Parameters affecting the dummy restraint system were studied; these parameters included the effect of no upper torso restraint, measurement of the amount of inertia-reel strap pullout before locking, measurement of dummy chest forward motion, and loads in the restraints. With the SI model, the dummies with no shoulder harness received head impacts above the concussive threshold.

Abstract: An aircraft capable of vertical short takeoff and landing is shown according to the teachings of the present invention as including four separate engine locations. In its most preferred form, the first and second engines are located on opposite sides of the fuselage of the aircraft and are pivotal about first and second axes which are in a plane perpendicular to the longitudinal axis of the aircraft and which are in front of the center of gravity of the aircraft. The third and fourth engines are located on opposite sides of the fuselage of the aircraft and are pivotal about third and fourth pivot axes which are parallel to the first and second axes of the first and second engines, respectively, but which are behind the center of gravity of the aircraft. The engines are mounted and rotated about their respective pivot axes by a system including a shaft having a first end operatively attached to the engine. The shaft is pivotally mounted to the fuselage of the aircraft by a bearing system including first and second tapered roller bearing members wedged between frusto-conical surfaces formed in a bearing mount, on the shaft, and on a collar removably attached to the shaft. A drive gear is attached to the shaft for rotation by a worm gear driven in turn by actuators such as hydraulic or electric motors. Thus, the engines at all four locations can be simultaneously pivoted about their respective axes.

Abstract: The Dryden Flight Research Facility has developed a unique research facility for conducting aerodynamic and fluid mechanics experiments in flight. A low aspect ratio fin, referred to as the flight test fixture (FTF), is mounted on the underside of the fuselage of an F-104G aircraft. The F-104/FTF facility is described, and the capabilities are discussed. The capabilities include (1) a large Mach number envelope (0.4 to 2.0), including the region through Mach 1.0; (2) the potential ability to test articles larger than those that can be tested in wind tunnels; (3) the large chord Reynolds number envelope (greater than 40 million); and (4) the ability to define small increments in friction drag between two test surfaces. Data are presented from experiments that demonstrate some of the capabilities of the FTF, including the shuttle thermal protection system airload tests, instrument development, and base drag studies. Proposed skin friction experiments and instrument evaluation studies are also discussed.

Abstract: Wind tunnel model tests support the hypothesis that a propeller tip vortex may subject a downstream wing surface to greater excitation than would be experienced by the aircraft fuselage side wall exposed to propeller-generated noise, ultimately transmitting this structural response to incident dynamic pressure to the cabin interior. Even if structure-borne excitations are less efficient than airborne excitations in the creation of cabin noise, the higher level of the former could still govern cabin noise levels.

Abstract: A composite spar wing assembly for use in model aircraft and/or air recreational vehicle construction. The composite spar wing structure is comprised of a composite spar having an I cross-sectional configuration which is adapted for snap engagement with spaced-apart transverse ribs therealong so as to form a wing structure. The composite spar is configured to have an elongate central web portion provided with upper and lower hollow flange members. The composite wing structures are connected to or through an aircraft fuselage by use of elongate spaced-apart connector elements which telescope respectively into the upper and lower hollow flange portions of the composie spar. A fuselage blade box assembly is selectively transversely provided through the fuselage and is provided with a segment of the composite spar so as to selectively receive the connector elements as previously described. A spring lock assembly is selectively provided in association with the aforementioned fuselage blade box assembly and the wing spar elements. The spring lock is configured to lockably engage the upper and lower connector elements inserted into the hollow flange members in contact therewith so as to prevent removal of the connector elements from engagement with the hollow flange portions of the composite spar section mounted in the fuselage blade box assembly and the wing structure. The composite wing spar can be selectively separated into two T-shaped components which are reassembled to form a longitudinally tapered modified wing spar having hollow flanges for use in forming a composite foam core wing structure.

Abstract: Operational and design features of twin-fuselage aircraft are outlined, noting capabilities of transporting 100-400 passengers at subsonic speeds at an efficiency of around 190 passenger mi/gal. Wings for two body aircraft are lighter and are designed more from an aerodynamics point of view due to reductions in the bending moment. A 280 passenger configuration would need a 172 ft wingspan, compared to a 155 ft wingspan for a conventional aircraft, but the conventional wings would have a larger area. The higher aspect ratio contributed to the increased efficiency of the twin body operation. A lower wetted fuselage area is calculated for the two body aircraft with passenger capacities over 190, and twin fuselages are shown to have a higher passenger packaging density than double-deck widebodies. Finally, simple compounding of existing aircraft such as the DC-9 into a two-body shape is projected to offer a 1.9 factor increase in passenger mi/gal.

Abstract: A pair of vanes (26, 28) extend laterally outwardly from the fuselage (10) of a swept wing aircraft immediately forwardly of the wings (12, 14). The preferred vanes (26, 28) have a trapezoidal plan form, an outwardly tapering thickness and an airfoil cross-sectional configuration. The vanes (26, 28) are substantially aligned with streamlines (46) during level flight. In a maneuver the vanes generate vortices which flow rearwardly over the upper surfaces of the inboard portions of the wings (12, 14), causing the inboard portions of the wings (12, 14) to stall at lower angles of attack than that of the wings (12, 14) alone would otherwise dictate and a raising of the downwash of the wings in the region of the tail. This lifting of the downwash results in the horizontal stabilizers (16, 18) of the tail being effective to create moments in opposition to "pitchup" of the aircraft.

Abstract: An analytical method is described for prediction of the interior noise levels for propeller-driven aircraft, given the exterior noise signature, its harmonic spectrum, and a description of the fuselage sidewall structure and various candidate "add-on" noise-control elements. The structural response is described by the theory of Koval, but simplified to consider the stiffeners as "smeared" elements. The incremental transmission loss (TL) due to add-on noise-control elements is derived from the Beranek and Work method. Comparisons between experimental data and the theory are presented. The method is reasonably accurate below the ring frequency, but is somewhat conservative at the normal incidence angle. This method is, however, expedient computationally, is economical, and permits rapid comparisons of noise-control penalties for various treatment concepts.

Abstract: The reingestion of hot exhaust gas can seriously reduce the performance of V/STOL aircraft. Past research on hot-gas ingestion is reviewed and design concepts that can be used to minimize ingestion are identified. Both the near field effects of the fountain flows created by multiple-jet configurations and the far field effects of wind or aircraft forward motion are considered. Techniques for minimizing hot-gas ingestion that are discussed include jet exit arrangements to simplify the fountain flow, shields designed to favorably redirect the fountain flow that impinges on the aircraft, minimizing the amount of hot gas projected ahead of the aircraft, and locating the inlets as high as possible.

Abstract: The fuselage (cabin) (15) of a propeller (45) driven aircraft (10) is isolated from vibration associated with the wake (46) of the propeller by vibration isolators (50, 55 and 110) in the wing and tail surfaces of the aircraft.

Abstract: Two baseline aircraft are configured in this study. The multibody aircraft has two fuselage bodies located at approximately 28% wing semispan. The single-body aircraft is configured in such a way as to provide a reference base with which the multibody aircraft can be compared. The aircraft are sized to provide the lowest direct operating cost configuration when transporting 771,620 lbs. over a distance of 3,500 nautical miles at a cruise speed of Mach 0.80. The aircraft are assumed to operate in the years 1990-1995, thus allowing for the incorporation of those technologies expected to be mature and available for production usage in 1985. In comparison with the single-body aircraft, the two-body aircraft shows reductions of 8.9% in wing weight, 7.7% in structural weight, 13.5% in block fuel weight, and 11.3% in direct operating cost.

Abstract: A structural performance and resizing finite element thermal analysis computer program was used in the reentry heat transfer analysis of the space shuttle. Two typical wing cross sections and a midfuselage cross section were selected for the analysis. The surface heat inputs to the thermal models were obtained from aerodynamic heating analyses, which assumed a purely turbulent boundary layer, a purely laminar boundary layer, separated flow, and transition from laminar to turbulent flow. The effect of internal radiation was found to be quite significant. With the effect of the internal radiation considered, the wing lower skin temperature became about 39 C (70 F) lower. The results were compared with fight data for space transportation system, trajectory 1. The calculated and measured temperatures compared well for the wing if laminar flow was assumed for the lower surface and bay one upper surface and if separated flow was assumed for the upper surfaces of bays other than bay one. For the fuselage, good agreement between the calculated and measured data was obtained if laminar flow was assumed for the bottom surface. The structural temperatures were found to reach their peak values shortly before touchdown. In addition, the finite element solutions were compared with those obtained from the conventional finite difference solutions.

Abstract: An aircraft having a plurality of fuselage members, one of which is provided with a replaceable and interchangeable payload carrying section. The replaceable section may be adapted to contain various equipment, ordinance and the like to satisfy different aircraft mission requirements. The replaceable and interchangeable sections are adapted to become an integral part of the fuselage section to which it is attached and permits the payload to be located in proximity to the aircraft's center of gravity such that variations in the equipment that constitutes the payload or discharge of all or part of the payload from the replaceable fuselage section will not substantially alter the aircraft's center of gravity.

Abstract: Acoustic data for the advanced-design SR-3 propeller at Mach numbers to 08 and helical tip Mach numbers to 114 are presented Several advanced-design propellers, previously tested in wind tunnels at the Lewis Research Center, are being tested in flight at the Dryden Flight Research Facility The flight-test propellers are mounted on a pylon on the top of the fuselage of a JetStar airplane Instrumentation provides near-field acoustic data for the SR-3 Acoustic data for the SR-3 propeller at Mach numbers up to 08, for propeller helical tip Mach numbers up to 114, and comparison of wind tunnel to flight data are included Flowfield profiles measured in the area adjacent to the propeller are also included

Abstract: An analytic study was performed to define the acoustical treatment weight penalties that are required to provide an interior noise level of 80 dBA in propfan-power ed aircraft at Mach 0.8 cruise. The prediction method, described in a companion paper, combines Koval's theory for cylindrical shell noise transmission loss (TL) with Beranek and Work's method for multilayered acoustic treatment analyses. Three fuselage diameters are studied which represent commuter, narrow-body, and wide-body aircraft. The calculated acoustic treatment weight penalties range from 1.7 to 2.4% of aircraft takeoff gross weight (TOGW) for add-on designs. Advanced noise reduction designs, those which permit structural modifications, reduce the acoustic treatment weight penalties to 1.5% TOGW for aluminum aircraft and from 0.74 to 1.4% TOGW for composite fuselage construction. The wide-body results agree with the weight penalty estimates of an earlier turboprop aircraft study.