Abstract: Modifications to a three-dimensional, implicit, upwind, unstructured-grid Euler code for aeroelastic analysis of complete aircraft configurations are described. The modifications involve the addition of the structural equations of motion for their simultaneous time integration with the governing flow equations. The paper presents a detailed description of the time-marching aeroelastic procedures and presents comparisons with experimental data to provide an assessment of the capability. Flutter results are shown for an isolated 45-deg swept-back wing and a supersonic transport configuration with a fuselage, clipped delta wing, and two identical rearward-mounted engine nacelles. Comparisons are made between computed and experimental flutter characteristics to assess the accuracy of the aeroelastic results.

Abstract: In a vertical take-off and landing vehicle, the invention is a method for improving stability by virtue of establishing conversion of thrust vectors over the center of gravity of the vehicle while in the hover position. Through the use of selected angles of inclination of the thrust-generating devices, a positive static stability of the aircraft is maintained. In addition, the spars supporting the thrust-generating devices are mounted in a fixed angular relationship to the centerline of the aircraft's fuselage, to achieve the desired inclination of the thrust vectors of the lift-generating devices toward the centerline of the vehicle, by simple rotation of the spars on which the lift-generating devices are mounted.

Abstract: An investigation has been conducted to determine a set of optimal design parameters for a single-stage-to-orbit reentry vehicle. Several configuration geometry parameters which had a large impact on the entry vehicle flying characteristics were selected as design variables: the fuselage fineness ratio, the nose to body length ratio, the nose camber value, the wing planform area scale factor, and the wing location. The optimal geometry parameter values were chosen using a response surface methodology (RSM) technique which allowed for a minimum dry weight configuration design that met a set of aerodynamic performance constraints on the landing speed, and on the subsonic, supersonic, and hypersonic trim and stability levels. The RSM technique utilized, specifically the central composite design method, is presented, along with the general vehicle conceptual design process. Results are presented for an optimized configuration along with several design trade cases.

Abstract: A method and apparatus for manufacturing panels and major airplane fuselage sections, including a reconfigurable fixture that holds panels for routing and drilling by accurate numerically controlled machine tools using original numerical part definition records, utilizing spatial relationships between key features of detail parts or subassemblies as represented by coordination holes drilled into the parts and subassemblies and making the parts and subassemblies intrinsically determinate of the dimensions and contour of the assembly.

Abstract: A lateral force device for displacing a towed underwater acoustic cable piding displacement in the horizontal and vertical directions having a spool and a rotationally mounted winged fuselage. The hollow spool is mounted on a cable with cable elements passing therethrough. The winged fuselage is made with the top half relatively positively buoyant and the bottom half relatively negatively buoyant. The winged fuselage is mounted about the hollow spool with clearance to allow rotation of the fuselage. The difference in buoyancy between the upper and lower fuselage maintains the device in the correct operating position. The wings are angled to provide lift in the desired direction as the fuselage is towed through the water.

Abstract: The fatigue growth of multiple cracks, of arbitrary lengths, emanating from a row of fastener holes in a bonded, riveted, lap joint in a pressurized aircraft fuselage is studied. The effects of residual stresses due to a rivet misfit, and of plastic deformation near the hole, are included. A Schwartz-Neumann alternating method which uses the analytical solution for a row of multiple colinear cracks in an infinite sheet (the crack-faces being subject to arbitrary tractions), is developed to analyze this MSD problem on a personal computer. It is found that for a range of crack lengths, a phenomena wherein the shorter cracks may grow faster than longer cracks may exist.

Abstract: The paper describes the first assessment study of a waverider-derived Mach 5 aircraft design using fuselage integrated-underslung over/under turboramjets with endothermic fuel. The study is based on a tanker-to-tanker mission, which begins at Mach 0.8 and 30,000 feet, with the vehicle accelerating to Mach 12 at constant altitude and then to Mach 5 while climbing to about 90,000 feet. The paper describes the vehicle, the aerodynamic analysis, and the propulsion system and its installed performance, structure design, and analysis. The mission simulation was run using the CSOTAV code developed by the NASA Langley Research Center.

Abstract: An aircraft includes a cockpit permanently connected to a fuselage base. At least one fuselage segment is detachably mounted to the support. In case of an imminent crash the segment is detachable from the remainder of the aircraft and can float to safety after being detached by having at least one parachute and an airbag secured to the segment.

Abstract: This invention outlines excitation means to transform the linear momentum of an underwing energized jet into rotational form in a selective manner to provide an asymmetric shear layer to increase compression wave reflection from the forward undersurface of a supersonic wing The wing compression energy is thereby recovered into useful work as an increase in pressure on the upward reflexed wing backside The upper surface of the shear layer is comprised of an array of vortices whose rotation is opposite to the wing circulation, providing the required angular momentum reaction The upper wing surface is flat to avoid generation of waves and an adverse angular momentum reaction above the wing The vortices below the wing are compressed by the underwing pressure, comprising a pressure shield to enhance the reflection The shear layer/vortex array grows in the stream direction due to augmented mixing with the underwing gap flow, which is turned and deflected upwards to provide a further increase in pressure on the upwards reflexed wing backside Fuselage bow shock energy is also recovered into useful work by a forward ring reflecting the conical shock inwards onto a suitably inclined shoulder An extendable nose spike allows the ring to intercept the conical bow shock at off-design Mach numbers The system in principle obviates wave drag to provide shock-free supersonic flight with improved efficiency and no sonic boom

Abstract: The present invention pertains to an aircraft that is capable of converting between vertical flight or helicopter mode flight, and horizontal flight or airplane mode flight where a two-bladed rotor is employed as both helicopter rotor blades in vertical flight and as a fixed wing in horizontal flight. In vertical flight, a bearing connection between two fuselage sections enables a forward section supporting the rotor blades to rotate relative to an aft section of the aircraft fuselage about the longitudinal axis of the aircraft. The exhaust or thrust force created by the mode of power (either a propeller engine or a turbine jet engine) is partially routed over the exterior of the aircraft to provide both vertical and horizontal thrust force, and in one embodiment a portion of the exhaust is routed through the interiors of the rotor blades and out exhaust ports at the blades' distal ends to rotate the blades in vertical flight and to provide a thrust force for the blades in horizontal flight. A control system controls the pitch position of the rotor blades to convert from vertical flight to horizontal flight by rotating the blades about their common lateral axis through opposite angles of substantially 90°.

Abstract: A sliding door for covering an opening (102) in an aircraft fuselage (100) including a sliding door member (10) mounted on the fuselage and adapted for sliding in the longitudinal direction of the fuselage between an opened and a closed position. Supporting and guiding means (30, 31, 32, 33) on the sliding door member (10) and the fuselage for supporting the sliding door in the opened and closed position and supporting and guiding sliding movement of the sliding door member (10) between the open and closed position.

Abstract: The influence of propulsion system variations and elastic fuselage behavior on the flight control system of an airbreathing hypersonic vehicle is investigated. Thrust vector magnitude and direction changes due to angle of attack variations affect the pitching moment. Low structural vibration frequencies may occur close to the rigid body modes influencing the angle of attack and lead to possible cross coupling. These effects are modeled as uncertainties in the context of a robust control study of a hypersonic vehicle model accelerating through Mach 8 using H-infinity and mu synthesis techniques. Various levels of uncertainty are introduced into the system. Both individual and simultaneous appearance of uncertainty are considered. The results indicate that the chosen design technique is suitable for this kind of problem provided that a fairly good knowledge of the effects mentioned above is available. The order of the designed controller is reduced but robust performance is lost which shows the need for fixed order design techniques.

Abstract: Reliable analytical methods that predict the structural integrity and residual strength of aircraft fuselage structures containing cracks are needed to help to understand the behavior of pressurized stiffened shells with damage, so that it becomes possible to determine the safe life of such a shell. Of special importance is the ability to determine under what conditions local failure, once initiated, will propagate. In this paper we shall present a reliable and efficient method for computing the energy release rate for cracks of varying length in a typical stiffened metallic fuselage under general loading conditions. The models used in the simulation are derived from an extensive analysis of a fuselage barrel section subjected to operational flight loads. Energy release rates are computed as a function of the length of the crack, its location, and the crack propagation mode.

Abstract: The development of an active aeroservoelastic missile fin using directionally attached piezoelectric (DAP) actuator elements is detailed. Several different types of actuator elements are examined, including piezoelectric polymers, piezoelectric fiber composites, and conventionally attached piezoelectric (CAP) and DAP elements. These actuator elements are bonded to the substrate of a torque plate. The root of the torque plate is attached to a fuselage hard point or folding pivot. The tip of the plate is bonded to an aerodynamic shell which undergoes a pitch change as the plate twists. The design procedures used on the plate are discussed. A comparison of the various actuator element shows that DAP elements provide the highest deflections with the highest torsional stiffness. A torque plate was constructed from 0.2032 mm thick DAP elements bonded to a 0.127 mm thick AISI 1010 steel substrate. The torque plate produced static twist deflections in excess of +/- 3 deg. An aerodynamic shell with a modified NACA 0012 profile was added to the torque plate. This fin was tested in a wind tunnel at speeds up to 50 ms/sec. The static deflection of the fin was predicted to within 6 percent of the experimental data.

Abstract: An approach to aerodynamic configuration optimization is presented for the high-speed civil transport (HSCT). A method to parameterize the wing shape, fuselage shape and nacelle placement is described. Variable-complexity design strategies are used to combine conceptual and preliminary-level design approaches, both to preserve interdisciplinary design influences and to reduce computational expense. Conceptual-design-level (approximate) methods are used to estimate aircraft weight, supersonic wave drag and drag due to lift, and landing angle of attack. The drag due to lift, wave drag and landing angle of attack are also evaluated using more detailed, preliminary-design-level techniques. New, approximate methods for estimating supersonic wave drag and drag due to lift are described. The methodology is applied to the minimization of the gross weight of an HSCT that flies at Mach 2.4 with a range of 5500 n.mi. Results are presented for wing planform shape optimization and for combined wing and fuselage optimization with nacelle placement. Case studies include both all-metal wings and advanced composite wings.

Abstract: The paper provides an analytical assessment of the flutter character of an unclassified National Aerospace Plane configuration known as the demonstrator. Linear subsonic, supersonic, and hypersonic analysis indicate that the vehicle is prone to body-freedom flutter resulting from the decrease in vibration frequency of the all-moveable wing at high flight dynamic pressures. As the wing-pivot frequency decreases, it couples with the vehicle short-period mode resulting in dynamic instability. A similar instability sometimes occurs when the pivot mode couples with the fuselage-bending mode. Also assessed, for supersonic flight conditions, are configuration variations that include relocation of the wing further aft on the lifting-body fuselage, and the addition of body flaps to the rear of the vehicle. These changes are destabilizing because they result in severe wing-pivot/fuselage-bending instabilities at dynamic pressures lower than the instabilities indicated for the original demonstrator. Finally, a two-point wing support and actuation system concept is proposed for the National Aerospace Plane, which if developed may (according to cursory analysis) enhance overall stability.

Abstract: The spacecraft comprises a fuselage (1), a wing (2), a power unit incorporating two liquid-propellant launching rocket engines (3), two liquid-propellant boost rocket engines (4), six transverse-thrust rocket engines (5) located in the spacecraft fuselage (1) on a rotatable ring (6), solid-propellant emergency deceleration rocket engines (7), and solid-propellant additional boosting rocket engines (8), a payload compartment (9), a crew compartment (10), a tail unit with two vertical fin struts (11) a bottom tailplane (12), and a top tailplane (13). The fuselage (1) is provided with a movable center conical body (14). The spacecraft landing gear has a swivelling tail wheel (21). The crew compartment (10) is interposed between the fin struts (11) under the top tailplane (13). The spacecraft is provided with an orbital maneuvering system whose final control elements are in fact low-thrust rocket engines (22) and (23), and gyrodynes. The crew escape system comprises an escape module (24) which is in fact a recoverable ballistic capsule held to the end face of the tail portion of the fuselage (1). The capsule (24) communicates with the crew compartment (10) through a tunnel (25) provided with means for crew transfer from the crew compartment (10) to the capsule (24). The capsule (24) has a front and a rear hatch, an aerodynamic decelerator, and a parachute system. The front portion of the capsule (24) has a heat-protective coating (36). The center of mass of the capsule is displaced towards its front portion.

Abstract: The flowfield generated by a helicopter in flight is extemely complex, and it has been recognized that interactions among components can significantly affect helicopter performance. In the present work a simplified model for the interaction of a rotor tip vortex with the helicopter fuselage is developed. The tip vortex is idealized as a single three-dimensional vortex tube, and the fuselage is moleded as an infinite circular cylinder

Abstract: The VTOL aircraft (10) includes a free wing (16) having wings on opposite sides of the fuselage (12) connected to one another for joint free rotation and for differential pitch settings under pilot, computer or remote control. On vertical launch, pitch, yaw and roll control is effected by the elevators (26), rudder (24) and the differential pitch settings of the wings, respectively. At launch, the elevator (26) pitches the fuselage (12) nose downwardly to alter the thrust vector and provide horizontal speed to the aircraft whereby the free wing (16) rotates relative to the fuselage (12) into a generally horizontal orientation to provide lift during horizontal flight. Transition from horizontal to vertical flight is achieved by the reverse process and the aircraft may be gently recovered in or on a resilient surface such as a net (66).

Abstract: Due to distortion of airflow streamlines, flow velocities and droplet size distributions measured around a moving aircraft can differ from freestream conditions. This can complicate measurements made from aircraft platforms. Potential flow calculations were used to predict airflow characteristics and the spatial distribution of different-sized droplets around the Lockheed Electra L-188 and Beechcraft King Air-200 aircraft at a variety of instrument mounting locations. Large deviations from freestream conditions were found to occur at certain locations on both aircraft near the fuselage and in regions of strong curvature. The number concentration of droplets 100–200 µm in diameter is most seriously affected by flow distortion effects. Calculation results were in reasonable agreement with measurements at a forward mounting location on the King Air.

Abstract: A tethered model gyroglider having a fuselage, a rotor support extending upwardly from the fuselage, and a unitary two-bladed rotor rotatably mounted onto the rotor support. A teetering mechanism permits the rotor to pivot about an axis proximate the center of the rotor. When the gyroglider is subjected to a relative wind generally from a forward direction, the forward moving blade of the rotor is raised and the rearward moving blade is lowered. This equalizes the aerodynamic forces about the longitudinal axis of the fuselage, thereby eliminating the tendency of the gyroglider to roll in the direction of the rearward moving blade.

Abstract: Three-dimensional thin-layer Navier-Stokes computations are presented for the F/A-18 configuration. The modeled configuration includes an accurate surface representation of the fuselage, leading-edge extension (LEX), and wing, both with and without leading-edge flap deflection. A multiblock structured grid strategy is employed to decompose the computational flowfield domain around the subject configuration. Steady-state solutions are obtained from an algorithm that solves the compressible Navier-Stokes equations with an upwind-biased, fluxdifference splitting approach. The results presented are based on a fully turbulent flow assumption, simulating the high Reynolds number flow conditions that correspond to a recent F/A-18 flight experiment. Good agreements between the computations and the flight test results are obtained for both surface flow patterns as well as surface pressure distributions. Furthermore, a correlation between the computed LEX vortex-core and the flight test results, observed by way of smoke visualization, is also presented.

Abstract: In a VTOL or STOL aircraft in which a lift device is housed internally within the aircraft fuselage or wing an outlet nozzle is stowable within the lift device exit aperture. A pair of doors, which when closed conceal the aperture from the sides of the nozzle, and a third folding member form the forward side of the nozzle and support a frame carrying variable vanes which vector the lift thrust.

Abstract: Aerodynamic and aerothermodynamic comparisons between flight and ground test for the Space Shuttle at hypersonic speeds are discussed. All of the comparisons are taken from papers published by researchers active in the Space Shuttle program. The aerodynamic comparisons include stability and control derivatives, center-of-pressure location, and reaction control jet interaction. Comparisons are also discussed for various forms of heating, including catalytic, boundary layer, top centerline, side fuselage, OMS pod, wing leading edge, and shock interaction. The jet interaction and center-of-pressure location flight values exceeded not only the predictions but also the uncertainties of the predictions. Predictions were significantly exceeded for the heating caused by the vortex impingement on the OMS pods and for heating caused by the wing leading-edge shock interaction.

Abstract: A device and method is disclosed whereby a protective cover is positioned on the upper surface of an aircraft wing. The cover comprises a lightweight, ultraviolet stabilized material which is resistant to propagation of tears and does not absorb water. Cover material (8), overhanging the leading or trailing edges of the wing, is locally cutout (13) in the area of protuberances to assist in providing a form fit when a plurality of cinch straps (22), attached to the leading and trailing edges, are tightened under the wing. Additional straps may be employed to secure the cover to the fuselage. The cover, when tightened, prevents significant ingress of air, protecting the upper surface.

Abstract: This invention is directed to stopped rotor Flipped Airfoil X-wing (FAX-WING.sup.™) aircraft which comprises: (a) A rotary wing flight mode which operates similar to a helicopter, wherein all main rotor airfoils rotate with leading edges into the oncoming airstream (neglecting forward motion) to provide lift, and utilize an anti-torque rotor to cancel main rotor torque. The fixed wing flight mode, including supersonic flight, utilizes all stationary main rotor airfoils for primary lift, such that all airfoil leading edges are positioned forward, meeting the oncoming airstream generated by forward aircraft motion. Two airfoils are forward swept 45 degrees, and the other two airfoils are aft swept 45 degrees. The transition mode for converting from rotary wing to fixed wing flight, and vise-versa, causes two adjacent airfoils to flip 180 degrees (in approximately 1/16 second) about their pitch axis, such that all airfoils have leading edges in the correct orientation for a particular flight mode. (b) Rotating flywheel inertia mass(es) capable of being coupled to the main rotor airfoils, via a flywheel clutch, for the purpose of rapidly starting or stopping airfoil angular rotation (in approximately 1 second) without applying adverse torque to the aircraft fuselage, and a spin up/spin down mechanism for applying angular momentum to the flywheel inertia mass(es) at gradual rates such that the anti-torque rotor and/or rudder can cancel torque. (c) Computer based flight control system capable of directing rotary wing, fixed wing, and transition therebetween. (d) Variable pitch ducted fan to provide forward thrust, coupled to common turboshaft engine(s), that also provide power to the main rotor.

Abstract: An elongated, rigid boom for the in-flight refueling of rotary wing aircraft in which a forward end of the boom is adapted for attachment to the fuselage of a rotary wing tanker aircraft and is of sufficient length to extend rearwardly of the tanker aircraft for the rear end of the boom to be clear of the tanker aircraft rotor path. The forward end of a funnel refueling drogue configured to receive the fueling probe of an aircraft to be refueled is swivelly attached to the rear end of the rigid boom and a fuel line supported by the boom extends from a connection into the tanker aircraft refueling tanks to a female refueling aircraft probe connection in the drogue. Pressurized air flowing within the boom is discharged from a downwardly directed nozzle and selectively from outwardly facing nozzles on each side of the boom adjacent the boom rear end with the nozzles being configured to establish a volume rate discharge as creates a vertical lifting force on the boom compensating for gravity and rotor downwash and boom side forces selectively directed horizontally outwardly in either direction for establishing yaw control of the boom.