Abstract: A large transport aircraft, for more than 400 passengers, with a first wing (3) extending from an intermediate point on the fuselage (1) and a second wing (4) extending from the rear of the fuselage, in proximity to the fin (2) The first wing comprises two halfwings (3a, 3b) swept backward, while the two halfwings (4a, 4b) of the second wing (4) are swept forward The first and second wings lie on two vertically spaced planes and are connected to one another by respective aerodynamic surfaces (5a, 5b) rigid in their plane in correspondence with the respective halfwings lying on the same side of the fuselage The first wing (3) can be either at a higher level than the second wing (4) or at a lower level

Abstract: The reduction of helicopter vibration is of great importance to the helicopter industry and considerable effort has been devoted to research into active methods for vibration reduction. The article describes how active control techniques can offer good vibration reduction over significant areas of the fuselage and also provide the capability to adapt to changing speed, rotor conditions and structural dynamics. There are three main areas on the helicopter where active vibration control techniques can be applied: at the rotor; at the main gearbox to fuselage interface, and within the fuselage itself. The last area is termed the active control of structural response (ACSR) and is considered in more detail. Finally, the performance of the ACSR technique in minimising vibration in helicopter structures is analysed using results which demonstrate the performance potential under differing conditions.

Abstract: A VTOL/STOL free wing aircraft (100) includes a free wing (110) having wings on opposite sides of a fuselage (102) connected to one another respectively adjacent fixed wing inboard or center root sections (117) fixedly attached to the fuselage (102) for free rotation about a spanwise axis (112). Horizontal and vertical tail surfaces (138, 140) are located at the rear end of a boom assembly (120) rotatably connected to the fuselage (102). A gearing (150) or screw rod (160) arrangement controlled by the pilot or remote control operator selectively relatively pivots the fuselage (102) in relation to the tail boom assembly (120) to enable the fuselage to assume a tilted or nose up configuration to enable VTOL/STOL flight.

Abstract: A circular wing aircraft in the form of a helicopter comprising a fuselage and a circular wing assembly. A structure is for mounting the circular wing assembly above the fuselage in a stationary manner. An air impeller unit is rotatively carried within the circular wing assembly. A device is for driving the air impeller unit to rotate about a central axis within the circular wing assembly, so as to provide lift and flight movement while yaw control is maintained.

Abstract: : Concerns over the safety of the continued use of aging transport aircraft have been voiced in the industry. A key component in the structural integrity of aging aircraft is the damage tolerance of fuselage structural repairs. The investigation focuses on the analysis and testing of contemporary repair methods and develops two improvements. The first technique, known as soft patching is appropriate for damage tolerant riveted repairs to incidental fuselage damage. Soft patching involves the use of high strength, moderate elastic modulus GLARE 3 fiber metal laminate patches. They extend the fatigue life of riveted repairs to monolithic aluminum fuselages while reducing life cycle costs. The second technique involves bonded crack patching of intact fatigue cracks in fuselage skins. An easy to use analysis program is presented along with analytical and test results of a low cost, high performance patch material known as GLARE 2. The findings demonstrate that GLARE 2 can replace the expensive boron/epoxy composites used in fuselage crack patching applications. Repair, Fatigue, Damage tolerance, Fuselage, Structures, Crack patching, Adhesive bonding, Surface pretreatment, Riveting, Advanced composite materials, Fiber metal laminates, GLARE, coefficient of thermal expansion.

Abstract: A configuration of four IO-channel single frequency C/A code narrow correlator spacing receivers is used to position two aircraft with respect to each other. The platforms used are U.S. Navy P-3 Orion aircraft cruising at speeds of 300 knots. Two receivers are mounted on each aircraft, with one antenna mounted on the fuselage and the other some 20 m aft on the tail boom. On each aircraft, the double difference carrier phase ambiguities between the two receivers can be resolved in minutes using the length constraint between the two antennas. The fixed ambiguity solution accuracy is sufficient to detect length variations at the level of 1-2 cm between the two antennas due to fuselage deformation caused by temperature variations. The carrier phase ambiguities between the two aircraft navigating within a few km from each other are successfully resolved using ambiguity constraint equations provided by the quadruple receiver configuration. This yields a relative position vector accurate to the cm-level. This highly accurate solution is then used to assess the accuracy achievable with a carrier phase smoothing of the code technique operating on all four receivers. A sub-meter three-dimensional SEP accuracy is demonstrated. The effect of code multipath caused by surrounding reflecting surfaces is analysed.

Abstract: An unmanned aerial vehicle (UAV) (100) having a toroidal fuselage (120) and a rotor assembly (170) including counter-rotating rotors coaxially mounted with respect to the toroidal fuselage incorporates ancillary aerodynamic structures (18) having a cambered airfoil profile to provide a nose-down pitching moment to counteract the nose-up pitching moment generated by airflow over the toroidal fuselage during forward translational flight of the UAV. The ancillary aerodynamic structures are symmetrically mounted in combination with the lateral sides of the toroidal fuselage so that the centers of lift are located aftwardly of the fuselage axis of the toroidal fuselage in forward translational flight modes. In a first embodiment, the ancillary aerodynamic structures (18) are fixedly mounted in combination with the toroidal fuselage (10) at a predetermined angle of incidence. In a second embodiment, the ancillary aerodynamic structures (19) are rotatably mounted in combination with the toroidal fuselage (10') to provide variable incidence ancillary aerodynamic structures for the UAV.

Abstract: Apparatus and method for recovering and arresting an aircraft on a floating platform is disclosed. The aircraft has a fuselage, a wing, and a source of propulsion for propelling the aircraft in flight. During landing approach, an inflatable cushion attached to the fuselage is deployed below the fuselage via inflation. In the inflated condition, a fastening material disposed on a downward facing surface of the cushion is thereby adapted to adhesively contact the floating platform deck so that the forward motion of the aircraft is slowed and arrested. In the preferred embodiment, the fastening material is one of male or female VELCRO® which is adapted to mate with corresponding VELCRO® material covering the platform deck.

Abstract: Composite Optimization Software for Transport Aircraft Design Evaluation (COSTADE) is being developed as a tool to support design build teams in their efforts to develop cost effective and feasible commercial aircraft fuselage structure. This paper contains a description of the development efforts of several of the major contributors to the code. The COSTADE tool currently contains a fabrication cost algorithm, structural analysis modules, a weight estimation module, manufacturing tolerance and constraint modules, and optimization algorithm modules. An application of COSTADE directed at evaluating and optimizing the ATCAS keel sandwich structure is outlined. Composite Optimization Software for Transport Aircraft Design Evaluation (COSTADE) is being developed by Boeing Commercial Aircraft Group in coordination with a number of subcontractors including the University of Washington (UW), Sikorsky Aircraft, Northrop, DowIUnited Technologies and the Massachusetts Institute of Technology (MIT). COSTADE enables a design build team to quantify and explore the complex interactions associated with aircraft design criteria, fabrication cost, manufacturing constraints and structural weight. This initiative is part of the NASAIBoeing Advanced Technology Composite Aircraft Structures (ATCAS) program. The objective of the ATCAS program is to develop an integrated technology and demonstrate a confidence level that permits the cost and weight-effective use of advanced composite materials in commercial transport fuselage structures. This paper is the first in a session containing six papers detailing the present status, development efforts and applications performed by several of the major developers of COSTADE. The primary developers and their roles are shown in Table 1. Cost Structural Manufacturing Optimization Software Model Analysis Constraints Algorithms Frame-

Abstract: A thin-layer Navier-Stokes code and a panel method code are used to predict the flow over a generic helicopter fuselage. The computational results are compared with pressure data at four experimental conditions. Both methods produce results that agree with the experimental pressure data. However, separation patterns and other viscous flow features from the Navier-Stokes code solution are shown that cannot be easily modeled with the panel method.

Abstract: A self-propelled passenger lift vehicle for transporting a mobility impaired passenger between a first location and first elevation and a second location and second elevation against an aircraft fuselage adjacent an aircraft door is disclosed. The vehicle comprises in combination a wheeled frame, a generally horizontal platform movably mounted on said frame having a leading edge portion and a support portion for supporting the passenger, a drive motor connected to at least one of the wheels for moving said vehicle from said first location to said second location, a brake mechanism connected to the at least one wheel for controllably stopping and preventing wheel rotation, and a position sensor in the leading edge portion of the platform for sensing the presence of the aircraft fuselage. The position sensor automatically actuates the brake to lock the drive wheel upon sensing contact with the aircraft fuselage. This feature minimizes damage to the aircraft due to operator error.

Abstract: A combined flying machine comprises a fuselage (1) in the form of a centralhick wing with a vertical open tunnel (2), in which there is mounted a lifting rotor (4). Said machine is also provided with outboard wings and a tail unit (8). The machine is equipped with a landing device on an air cushion (19) that surrounds an outlet from the tunnel. A power plant of said flying machine comprises two engine modules (13), disposed from two sides of the tunnel (2) and connected with the lifting rotor (4) and the propulsion propellers (12). The area of the tunnel cross-section in the plane of the lifting rotor rotation amounts to 0.3 to 0.8 of the area of the landing device air cushion, which ensures a safe landing of the flying machine on an unprepared landing site even with a failure of one of the engine modules.

Abstract: A system for converting all or part of a passenger cabin (2) in a fuselage (1) of an aircraft for the carriage of freight, the cabin having internally finished longitudinal walls (10) and ceiling (8) defining an internal cross-sectional profile of substantially uniform cross section. The system includes a dismountable rigid liner structure having an external profile receivable within the internal profile of the cabin without contacting the walls and ceiling. It is formed of interlocking sets of parts (20, 26) individually small enough to pass through a cargo door (14) in the fuselage and is supported between ends of the cabin solely by releasable anchorages (32) engaging seat tracks (16) installed in the cabin, the liner structure including transverse cargo restraints (50) releasably secured thereto for restraining movement of cargo through said liner structure. The modules form a series of longitudinally connected compartments, each compartment including a truss element (26) and anchors for securing a cargo restraint at its front end, the rearmost compartment being closed at its rear end by a cargo restraint (60, 62) connected to the fuselage of the aircraft in place of a removable rear cabin bulkhead (4), which may be relocated ahead of the liner structure if only a part of the cabin is converted.

Abstract: The air intake includes two main flaps (1, 2) facing each other and each pivoting about a pin (4, 5) adjacent to a wall (6, 7) extending the flap in question (1, 2) rearwards, and in which a first boundary layer bleed (8a, 8b) is arranged, the pins (4, 5) being substantially parallel to the plane of the wings or of the fuselage so that the flaps (1, 2) move perpendicularly to this plane, and a ramp (11) pivoting about a pin (12) adjacent to its leading edge (11a) by which it is adjacent to the wings and to the fuselage and substantially parallel to the pins (4, 5) is upstream of the flap (1) which is closest to the wings and with which it delimits another boundary layer bleed (13) of variable cross section.

Abstract: I T is shown in Ref. 1 that the presence of a fuselage significantly promotes the breakdown of delta wing leadingedge vortices. Applying the equivalent angle-of-attack concept only accounted for a fraction of the measured effect of the fuselage, predicting a 17% increase of the effective angle of attack compared to the measured 45% increase. The present comment describes a flow mechanism that can explain this discrepancy, i.e., the wing-camber effect generated by the fuselage-induced upwash along the leading edge of the delta wing. The effect of longitudinal camber on the breakdown of the leading-edge vortex on a slender delta wing is large (Fig. 1). For the same maximum local angle of attack on the delta wing max a positive camber of Aa/amax = 1 delays breakdown to occur downstream of the trailing edge, whereas a negative camber of the same magnitude, Aa/amax = — 1, causes burst to occur very close to the apex. Obviously, for a pitching delta wing the pitch-rate-induced camber will have similarly large effects on the breakdown of leading-edge vortices. It is described in Refs. 4 and 5 how the roll-rate-induced camber effect would be very similar to the pitch-rate-induced camber effect. In both cases, it is the motion-induced change of the local angle of attack at the leading edge that matters (Fig. 2). For the maximum reduced frequency and amplitude used in the roll oscillation test of a 65-deg, sharp-edged delta wing, the roll-rate-induced camber at the trailing edge was AaLE/a = 0.31. Following the suggestion in Ref. 7, static tests were performed with models deformed to produce the roll-rateinduced camber (Fig. 3). The results were as expected; i.e., the twisted-up side of the delta wing experienced later vortex breakdown than the opposite, twisted-down side, approximately at 70% chord compared to 45% chord (for zero roll angle <£ = 0). It should be noted that the variation of the induced angle of attack along the leading edge is linear, proportional to the local semispan, in the examples given in Figs. 1-3, whereas the variation is nonlinear, roughly inversely proportional to the local semispan, in the case of the body-induced wing camber. However, the results in Ref. 7 serve to illustrate how large the effect of camber is on vortex breakdown. Based upon these results one can expect that the bodyinduced negative wing camber along the delta wing leading edge is a very important parameter. Thus, a better approach than using any mean-alpha value for the delta wing, such as the equivalent angle of attack, is to use the body angle of attack a together with the body-induced wing camber at the trailing edge AaLE/a, to correlate measurements of the breakdown of delta wing leading-edge vortices in the presence of a fuselage.

Abstract: Computational analysis of flow over the F/A-18 aircraft is presented along with complementary data from both flight and wind tunnel experiments. The computational results are based on the three-dimensional thin-layer Navier-Stokes formulation and are obtained from an accurate surface representation of the fuselage, leading-edge extension (LEX), and the wing geometry. However, the constraints imposed by either the flow solver and/or the complexity associated with the flow-field grid generation required certain geometrical approximations to be implemented in the present numerical model. In particular, such constraints inspired the removal of the empennage and the blocking (fairing) of the inlet face. The results are computed for three different free-stream flow conditions and compared with flight test data of surface pressure coefficients, surface tuft flow, and off-surface vortical flow characteristics that included breakdown phenomena. Excellent surface pressure coefficient correlations, both in terms of magnitude and overall trend, are obtained on the forebody throughout the range of flow conditions. Reasonable pressure agreement was obtained over the LEX; the general correlation tends to improve at higher angles of attack. The surface tuft flow and the off-surface vortex flow structures compared qualitatively well with the flight test results. To evaluate the computational results, a wind tunnel investigation was conducted to determine the effects of existing configurational differences between the flight vehicle and the numerical model on aerodynamic characteristics. In most cases, the geometrical approximations made to the numerical model had very little effect on overall aerodynamic characteristics.

Abstract: Nozzles (3, 5) for discharging high velocity air are positioned on either side of an aircraft engine exhaust duct (1) exit High velocity air is discharged tangential to an outer surface of each nozzle (3, 5) and follows the contour thereof The nozzles (3, 5) are positioned so that when the high velocity air separates from the outer surface of the respective nozzles (3, 5), it flows outwardly relative to the aircraft's fuselage structure The high velocity air from the nozzle (3) closest to the aircraft's fuselage structure impinges on the exhaust gas stream (35, 37), deflecting the stream away from the fuselage (43) The high velocity air from the nozzle (5) furthest from the fuselage creates a low pressure area which deflects the exhaust gas stream (35, 37) away from the fuselage In combination, the two nozzles (3, 5) deflect the exhaust gas stream (35, 37) away from the fuselage to a significant degree

Abstract: An issue of concern in aging aircraft is the growth of multiple cracks emanating from a row of fastener holes, typically in a pressurized aircraft fuselage lap splice. This multisite damage (MSB), or widespread fatigue damage, if allowed to progress, can suddenly become catastrophic. The understanding of the failure behavior dictates the level of compromise between safety and economy. The complexity of the structure due to various stiffening elements makes it unamenable to a simple direct analysis. A two-step elastic finite element fatigue analysis combining a conventional finite element method and the Schwartz-Neumann alternating method with analytical solutions is developed to understand fatigue growth of multiple cracks and to obtain a first estimate of the residual life of a stiffened fuselage shell structure with MSD in the riveted lap joint. The analysis procedure is validated by simulating a laboratory fatigue test on a lap joint in a flat coupon. Both the coupon and the shell panel are found to have fatigue lives only up to the first linkup of neighboring crack tips.

Abstract: A free wing aircraft including a free wing pivotally supported about a spanwise axis for flight in a free wing mode of operation with respect to the fuselage is disclosed. The free wing is capable of being deflected into a nose down configuration sufficient to create an aerodynamic braking effect to decrease air speed after landing upon roll-out. This aerodynamic braking effect can also be sufficient to create negative lift that plants the aircraft firmly down, increasing the coefficient of friction and allowing the aircraft brakes to be more effective relative to a fixed wing aircraft at touchdown.

Abstract: The present invention provides a desired design approach for retuning a rotorcraft airframe by utilizing resilient connections between major airframe components. In the preferred embodiment, the design employs resilient material, preferably elastomeric material, in the mounting system between the vertical tail and the tailboom of a helicopter as a method of providing dynamic detuning of the tailboom torsion/lateral mode. Other possibilities include similarly mounting the horizontal stabilator and the wing onto the airframe. Using elastomeric or similar devices to soft mount the vertical tail of a helicopter on the tailboom can efficiently and effectively alter the natural frequencies of the entire aircraft. By choosing the spring constant of this connection, the designer can place fuselage frequency at a desired value.

Abstract: The characteristics of widespread fatigue damage (WSFD) in fuselage riveted structure were established by detailed nondestructive and destructive examinations of fatigue damage contained in a full size fuselage test article. The objectives of this work were to establish an experimental data base for validating emerging WSFD analytical prediction methodology and to identify first order effects that contribute to fatigue crack initiation and growth. Detailed examinations were performed on a test panel containing four bays of a riveted lap splice joint. The panel was removed from a full scale fuselage test article after receiving 60,000 full pressurization cycles. The results of in situ examinations document the progression of fuselage skin fatigue crack growth through crack linkup. Detailed tear down examinations and fractography of the lap splice joint region revealed fatigue crack initiation sites, crack morphology and crack linkup geometry. From this large data base, distributions of crack size and locations are presented and discussions of operative damage mechanisms are offered.

Abstract: A T-38 trainer aircraft that is modified to have the wings of an F-5 fighter aircraft. The F-5 wings have leading edge flaps which increase aircraft maneuverability and allow the plane to be landed at a lower speed. The F-5 wing is coupled to existing wing joints of the fuselage by fittings which increase the angle of incidence of the aircraft approximately 2°. The increased angle of incidence provides the pilot with a greater over the cone field of view. The modified T-38 also has a pair of forward strakes located on the leading edge of the aircraft nacelles. The forward strakes further increase the performance of the plane. Attached to the tail of the aircraft are a pair of aft strakes which counteract the pitch created by the forward strakes.

Abstract: Tandem-engine aircraft propulsion module attachable to an aircraft fuselage and/or to the aircraft wings directly. This propulsion module contains fore-and-aft in-line engines and an intermediate fuel compartment. A subhousing of this module has engine instrumentation and controls placed for overhead introduction into an underlying aircraft fuselage cabin when the module is secured on top thereof. Propulsion modules of this invention are readily removable and replaceable for ease of inspection and maintenance, and if like-sized are mutually interchangeable. These modules are inherently economical and enable cost-effective fuselage and dry-wing construction. They also enhance flight safety, not only in comparison with single-engine aircraft but also especially as compared with off-axial twin-engine aircraft.

Abstract: A fighter aircraft achieves low aerodynamic drag and radar signature without sacrificing flight performance through a unique arrangement of the main weapons bay, the auxiliary weapons bays, and the main landing gear. Separate main and auxiliary weapons bays permit a narrower fuselage than could be obtained with a single common bay. Also, the auxiliary weapons bays and the landing gear can be aligned outboard of the main weapons bay without needing to increase the length or width of the aircraft. The air intake ducts extend aft from the intake and curve upwardly and inwardly over the main weapons bay. The result of the design configuration is an aircraft which has a forward aspect reduced to the minimum necessary to accommodate the components that need forward visibilities, which translates to minimum aerodynamic drag and radar signature.

Abstract: This paper describes the development of a technique for predicting the failure strength of fiber reinforced advanced composite structures which have damage induced cracks and flaws. The technique, described as a Damage Zone Model (DZM) simulates the manner in which the stress concentrations arising from such flaws affect the stable growth of damage and ultimately failure strength of the structure. The concept of a DZM is to describe the multiple damage modes which occur in laminated composite structures via a strain softening material law which permits controlled unloading of material at, for example, a crack tip as applied loads and the size of the crack increase. The technique has been implemented in a commercially available nonlinear finite element code and has been subjected to extensive experimental verification. Experimental fracture strength for a wide range of test specimen sizes and materials have been predicted to a degree of accuracy unobtainable with more conventional fracture mechanics techniques. As an analytical tool it may prove invaluable in designing lighter, safer composite products more quickly and with less testing.