Abstract: An unmanned arerial vehicle (UAV) has a toroidal fuselage and a rotor assembly having a pair of counter-rotating rotors secured in fixed coaxial combination with the toroidal fuselage to provide a vertical takeoff and landing (VTOL) capability for the UAV. One embodiment of the VTOL UAV is especially configured for ground surveillance missions by the inclusion of an externally mounted, remotely controllable stowable sensor subsystem (250) that provides an azimuthal scanning capability and a predetermined elevation/depression scanning capability to accomplish the ground surveillance mission and a foldable landing gear subsystem (300) to facilitate landing of the VTOL UAV at unprepared ground surveillance sites. The foldable landing gear subsystem includes a plurality of legs (302), one end of each leg being detachably secured in combination with the toroidal fuselage, a foot that includes a pad member (312) pivotally attached to the other end of each leg, and a non-structural hinge secured to each leg and the toroidal fuselage. The non-structural hinges (306) provide the capability to fold the landing gear subsystem to a stowage configuration wherein each leg, pivoting foot combination is folded within the envelope of the toroidal fuselage.

Abstract: An efficient methogology is presented for defining a class of airplane configurations. Inclusive in this definition are surface grids, volume grids, and grid sensitivity. A small set of design parameters and grid control parameters the process. The general airplane configuration has wing, fuselage, vertical tail, horizontal tail, and cannard components. The wing, tail, and canard components are manifested by solving a fourth-order partial differential equation subject to Dirichlet and Neumann boundary conditions, and the solution is expressed as a Fourier Series. The fuselage has circular cross section, and the radius is an algebraic function of four design parameters and an independent computational variable. Volume grids are obtained through an application of the Control Point Form method. Grid Sensivity is obtained by applying the automatic differentiation precompiler ADIFOR to software for the grid generation. The computed surface grids, volume grids, and grid sensivity are suitable for a wide range of Computational Fluid Dynamics simulation and configuration optimization.

Abstract: A supersonic airplane having four or two supersonic engines and one or more boost engines. During take-off and initial climb the supersonic engines are operated at a lower thrust setting within acceptable noise limits, and the subsonic engine(s) is operated to provide boost thrust to enable the airplane to operate a take-off and climb. During cruise, the subsonic engines are in a nonoperating mode, and the supersonic engines alone provide the thrust for supersonic operation. In one embodiment, one subsonic engine is deployed on one side of the fuselage during the operating mode. In a second embodiment, there are two subsonic engines deployed on opposite sides of the fuselage. In another embodiment, a single subsonic engine is installed inside the fuselage.

Abstract: A primary or backup flight control system for controlling the flight path of an aircraft or spacecraft using wireless transmitters (15 and 16) located near the flight deck (14) and positioned within, or on, the outer skin of the fuselage (10). These transmitters (15 and 16) send pilot-generated, wireless, e.g. infra-red, flight control signals (17 and 18) to receivers (19 and 21) located within, or on, the engines, wings, and tail assembly surfaces. These signals (17 and 18) are then directed to their respective engines or moveable flight control surfaces. These wireless control signals (17 and 18) are transmitted externally of the fuselage structure (10) and the transmission and receipt of these signals takes place on a single aircraft or spacecraft structure. In case of fuselage damage, e.g., due to structural failure, internal explosions, air-to-air collision, or military combat, which damages the primary flight control system (hydraulic, wire or cable), the present system will remain operative.

Abstract: The design of structurally efficient joints in aircraft fuselage structures and wing skin splices is addressed. It is contended that the joints should be designed first and the gaps in between filled in afterwards, taking pains not to optimize the basic structure first and then discover that it either cannot be assembled or that, when it is assembled, it is full of weak-link fuses. Both adhesively bonded and mechanically fastened joints are covered. Analogies are drawn between the characteristics of both classes of joints. The aspects of static joint strength and fatigue lives are included. The work is applicable to metallic as well as composite structures, and covers both high-load wing joints which have already been tested and new ideas for fuselage splices which have not. The effects of flaws and defects are associated with the need for damage tolerance, particularly in fuselage structures.

Abstract: The craft is for hovering flight, vertical takeoff and landing, and horizontal forward flight. It has a tail-sitting fuselage and a ducted fan mounted to the fuselage aft to provide propulsion in both (a) hovering and vertical flight and (b) horizontal forward flight. At each side is a floating wing, supported from the fuselage for passive rotation (or an actuator-controlled optimized emulation of such rotation) about a spanwise axis, to give lift in forward flight. The fuselage attitude varies between vertical in hovering and vertical flight, and generally horizontal in forward flight. Preferably the fuselage is not articulated; there is just one fan, the sole source of propulsion, rotating about only an axis parallel to the fuselage; and thrust-vectoring control vanes operate aft of the fan. Preferably at each side a small, nonrotating wing segment is fixed to the fuselage, and the floating wing defines--along its trailing portions--a corner notch or slot near the fuselage; forward portions of the fixed wing segment are within this notch. Preferably the spanwise axis is along a surface of the floating wing, and a long hinge supports that wing from the fixed wing segment, within the notch. During vertical and transitional flight characteristically the leading edge of the floating wing is down relative to the fuselage axis.

Abstract: Static pressure sensing and free airstream temperature sensing is obviated in a method for determining aircraft velocity relative to an airmass by utilizing a static pressure and temperature component present in pressure variations sensed by a total pressure sensor located on a rotating arm mounted on the fuselage of an aircraft. A harmonic analysis of a quasi-sinusoidal total pressure variation includes a determination of steady state and both first and second harmonic components. These harmonic components are used to determine the static pressure, the free airstream temperature and the aircraft airspeed without separately sensing these parameters.

Abstract: An aircraft control system for controlling an aircraft, particularly a free wing aircraft in low speed or hover regimes. An air speed sensor measures air speed of the aircraft and outputs an air speed signal to a control processor which processes the air speed signal with a speed control input signal. A control actuator actuates an aircraft control surface in response to the control surface control signal. The air speed sensor may include a shaft mounted impeller located in an airstream of the aircraft. A rotational speed sensor, coupled to the impeller, measures a rotational speed of the impeller and outputs a rotational speed signal as the air speed signal. In an alternative embodiment, the air speed sensor may include a vane located in an airstream of the aircraft and deflected in response to air flow in the airstream. In another embodiment, the speed sensor may include an angular position sensor which measures an angle between a free wing and the aircraft fuselage and outputs an angle measurement signal as the air speed signal. The aircraft control surface may comprises a control boom pivotally attached to a fuselage of the aircraft of a trim tab pivotally attached to a fuselage of the aircraft.

Abstract: An improved VTOL/STOL free wing aircraft providing damping and absorption of shock landing loads upon landing A pair of resilient struts is provided, projecting forwardly from the trailing edge of either side of the fuselage when the fuselage is tilted Preferably, the aircraft includes a pair of articulated tail booms, the strut being a portion of the tail boom extending forward from the pivot axis of the tail boom Landing wheels are disposed on the strut in tandem spaced relationship The resiliency of the strut causes the strut to act as a leaf spring and thus dampen shock landing loads Operatively secured to the bottom surface of the fixed wing portions and the forward portion of the landing gear struts is a pair of dashpots for absorbing the shock landing loads

Abstract: The invention relates to a helicopter fuselage of the type including a central structure to which are linked a front structure, a rear structure and a landing gear, and which supports a transmission gearbox, a main rotor and at least one engine (50), and of the type in which the central structure includes a skeleton (16) fitted with covering elements which define the external shape of the fuselage. According to the invention, the skeleton (16) of the central structure exhibits substantially the shape of a regular hexahedron consisting of skeleton panels (28, 30, 32, 34, 36, 38) assembled together. The invention finds an application especially for producing light helicopters.

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 model airplane flight simulator which enables a model airplane to be maneuvered by a remote control device through pitch, yaw and roll without movement over ground. The model airplane includes most of the attributes of a flying model such as one and preferably two, motor-driven propellers for generating airflow over control surfaces which include actuator controlled ailerons, elevator and rudder and a remote control device for selectively controlling these features from the remote location. The control device may utilize hard wiring or radio signal to activate the actuator system. No airfoil for wing lift is required in that the model airplane is supported within a rigid inner ring or hoop to which wing or fuselage extension rods coaxially extending from each wing tip or from each end of the fuselage are connected at diametrically opposing points on the inner ring for free rotation of the model airplane within the inner ring about one rotational axis. The inner ring is itself supportively connected for free rotation within a rigid outer ring or hoop along another rotational axis orthogonal to that passing through the wing or fuselage extensions. The outer ring, which may also be a half ring, is supportively connected for free rotation to a stationary support or base member about yet another rotational axis orthogonal to and intersecting the second axis at a mid-point between the bearing members connecting the inner and outer rings.

Abstract: An airplane fuselage panel including a sheet having peripheral edges routed on routing surfaces, while the sheet is held immobile on a fixture, using a routing end effector carried by a precision computer controlled robot that is directed to the routing surfaces using a digital dataset taken directly from digital engineering part definition records. The sheet has coordination holes drilled while on the fixture using a drilling end effector carried by the precision computer controlled robot that is directed to drilling locations using the digital dataset taken directly from the digital engineering part definition records to accurately locate the hole locations relative to the peripheral edges. The airplane fuselage panel also includes parts, including stringers, stringer clips and shear ties, each having coordination holes drilled by computer controlled drills at locations that will match with corresponding coordination holes in the sheet, so that the parts will be accurately located in positions called for in the digital engineering part definition records when the coordination holes in the parts and the coordination holes in the sheet are aligned and the parts are riveted to the sheet in the accurately located positions.

Abstract: A VTOL/STOL free wing aircraft includes a free wing having wings on opposite sides of a fuselage connected to one another respectively for free rotation about a spanwise access. Improved control upon landing of the aircraft is achieved by utilizing a variable pitch propulsion system, enabling the pitch of the propeller to be varied corresponding to the speed of the aircraft and angle of approach upon descent.

Abstract: Application of a wake roll-up method coupled with the vortex lattice method and approximate expressions for the receiver fuselage effect have been used to determine the induced loads on a Hercules receiver aircraft behind a KC10 tanker. The induced loads depend strongly on the vertical position of the receiver wing and fin relative to the tanker wing wake. In the case of steady sideslip there is a large decrease in the directional stability of the receiver as quantified by the gradient of the rudder angle versus sideslip. This is due mainly to the combined effects of the yawing moments due to bank, yaw and side displacements. Minimum directional stability corresponds to the tip of the receiver fin intersecting the tanker wing wake. The associated aileron angle is two to three times the value in free air in agreement with flight test data. Solution of the linearized equations of motion gives three lateral characteristic oscillations for the air-to-air refuelling case. These include the usual Dutch roll osc...

Abstract: Aircraft structural integrity is a major concern for airlines and airframe manufacturers. To remain economically competitive,airlines are looking at ways to retire older aircraft, not when some fixed number of flight hours or cycles has been reached, butwhen true structural need dictates. This philosophy is known as "retirement for cause." The need to extend the life of commer-cial aircraft has increased the desire to develop nondestructive evaluation (NDE) techniques capable of detecting critical flawssuch as disbonding and corrosion. These subsurface flaws are ofmajor concern in bonded lap joints. Disbonding in such ajointcan provide an avenue for moisture to enter the structure leading to corrosion. Significant material loss due to corrosion cansubstantially reduce the structural strength, load bearing capacity and ultimately reduce the life of the structure.Recent advances in technology have spawned a number of new NDE techniques that provide quantitative information aboutflaws in aircraft structures. In a effort to develop NDE techniques with improved reliability and reduced cost, several improve-ments in large area scanning techniques have been made. A few of these techniques currently being developed are thermalimaging, acoustic emission, scanning array ultrasonics, coherent optics, radiography and magnetic field visualization.One technique in particular that has proved quite promising is infrared thermal imaging. Thermography has a number ofadvantages as an inspection technique. It is a totally noncontacting, nondestructive, imaging technology capable of scanning alarge area of the aircraft fuselage for defects in a matter ofa few seconds. Advances in fast, inexpensive image processors haveaided in the attractiveness of thermography as a NDE technique. These image processors have increased the signal to noiseratio of thermography and facilitated significant advances in post-processing. The resulting digital images can also be archivedfor later comparison with the results of subsequent inspections thus providing a means of monitoring the evolution of damagein a particular structure.The National Aeronautics and Space Administration's Langley Research Center has developed a thermal NDE systemdesigned for application to disbonding and corrosion detection in aircraft skins. By injecting a small amount of heat into thefront surface of an aircraft skin, and recording the time history of the resulting surface temperature variations using an infraredcamera, quantitative images of both bond integrity and material loss due to corrosion can be produced. This paper will presenta discussion of the development of the thermal imaging system as well as the techniques used to analyze the resulting thermalimages. The analysis techniques presented represent a significant improvement in the information available over conventionalthermal imaging due to the inclusion of data from both the heating and cooling portion of the thermal cycle.Results of laboratory experiments on fabricated disbond and material loss samples will be presented to determine the limita-tions of the system. Additionally, the results of actual aircraft inspections will be shown, which help to establish the field appli-cability for this technique. A recent application of this technology to aircraft repairs using boron /

Abstract: : In October 1993 the FAA Technical Center conducted a vertical drop test of a narrow- body fuselage section. This test was structured to determine the impact response characteristics of some typical items of mass found aboard transport airplanes to assess the adequacy of the design standards and regulatory requirements for those components. A primary objective of this test was to determine the dynamic response characteristics of the onboard overhead stowage bins and auxiliary fuel tank system, as well as the fuselage section itself, when subjected to a potentially survivable impact. The dynamic impact environment and the resultant response of the onboard overhead stowage bins and auxiliary fuel tank system were characterized. The structural support reactions for those onboard items of mass were measured and compared to predicted values which were based on static analyses and tests. The test was intentionally structured to impose a dynamic load condition in excess of the current design and certification requirements for the onboard items of mass so that the dynamic fracture loads and modes of fracture for those components could also be determined and evaluated.

Abstract: It is now well established that the leading edge is a critical region for boundary layer transition to turbulence on a swept-back wing. Under the appropriate conditions, transition can occur on the attachment line itself (x=0) and this is particularly important since, in the absence of subsequent relaminarisation, full chord, turbulent flow on both upper and lower surfaces will be produced. It has been demonstrated in previous papers [1,2] that, for an impervious surface, the attachment-line transition process can be instigated at Reynolds numbers which are very much lower than those associated with linear (and non-linear) stability theory. Moreover, there is a very clearly defined lower boundary for the transition Reynolds number in the presence of large forcing disturbances. Large disturbances can take the form of surface roughness or they can emanate from the wing fuselage junction. Under these circumstances the transition process is commonly referred to as “attachment-line contamination” In the context of the laminar flow aircraft, attachment-line contamination is known to be a serious practical problem. Therefore, the determination of the effect of surface transpiration on attachment-line contamination

Abstract: In a tail unit for a commercial aircraft having a pressurized cabin, the trimmable elevator assembly and the rudder assembly are connected to a shortened fuselage tail section. The fuselage tail section is shortened because at least the upper contour line of the tail section and the two lateral contour lines of the tail section each have at least one curvature reversal point in the outer contour. Additionally, the center box (23) of the elevator assembly passes through the tail section (16) in the area of the tail ribs (20, 22) which support the rudder assembly. This shortening of the tail section allows better utilization of the cabin space in the aircraft. Additionally, it reduces the aerodynamic drag and the total weight of the tail unit.

Abstract: A canard wing surface for aircraft, especially supersonic aircraft, which can be lowered and retracted flush into the fuselage of the aircraft when no longer needed. The canard is hinged to an extendible platform at or near its upper edge, so that a raising mechanism on the platform can raise the canard from a position integral with the skin of the fuselage to a fully extended position as a curved wing. The platform is extended to allow the canard to clear the fuselage, and may be rotated to adjust the angle of attack of the canard from nearly flat, for the transition to high-speed flight, to a range of low-to-medium angles of attack when used as a lift surface during takeoff and landing, to nearly vertical for use as a speed brake during the landing roll. If desired, multiple canard surfaces can be used on an aircraft, arranged along the fuselage.

Abstract: Many air-breathing hypersonic vehicle design concepts utilize the lower surface of an elongated fuselage forebody to provide aerodynamic compression for a supersonic combustion ramjet module, or a scramjet. This highly integrated design approach creates the potential for an unprecedented form of aeroelastic-p ropulsive interaction in which deflections of the vehicle fuselage give rise to propulsive force and moment variations that may impact the vehicle's flight dynamic characteristics. This investigation examines the potential for such interactions using a math model that describes the longitudinal flight dynamics, propulsion system, and first seven elastic modes of a hypersonic vehicle concept. Estimates of the propulsion system sensitivity to angle-ofattack variations and modal fuselage deflections are presented and compared with predictions based on an earlier reference. The combined aeroelastic-propulsive model is used to illustrate a variation in the vehicle's longitudinal flight dynamics with the propulsion system sensitivities. Numerical values for the completed model at flight conditions of Mach 6 and Mach 10 are presented.

Abstract: As part of NASA's composite structures crash dynamics research, a general aviation aircraft with composite wing, fuselage and empennage (but with metal subfloor structure) was crash tested at the NASA Langley Research Center Impact Research Facility. The test was conducted to determine composite aircraft structural behavior for crash loading conditions and to provide a baseline for a similar aircraft test with a modified subfloor. Structural integrity and cabin volume were maintained. Lumbar loads for dummy occupants in energy absorbing seats wer substantially lower than those in standard aircraft seats; however, loads in the standard seats were much higher that those recorded under similar conditions for an all-metallic aircraft.

Abstract: Steady flowfields about a generic aircraft fuselage were simulated numerically by integration of the NavierStokes equations, including a two-equation (k-s) turbulence model. A steady, sonic, underexpanded jet issuing from a small square aperture in the fuselage surface was used to model the results of an internal explosion that ruptured the aircraft's skin. The computed solutions include simulations of a wind-tunnel test, and of flight at cruise conditions typical of a large transport aircraft. In each case, both the jet-on and jet-off situations were considered. Details of the computations are presented, and features of the flowfield are discussed. Comparisons were made with experimental data in terms of surface static pressure distributions and total pressure loss profiles, and found to be acceptable for engineering purposes.

Abstract: The suspension device comprises, for linking the base (5b) of the transmission gearbox to the fuselage (6), two parallel rods (14), articulated (15) onto lateral supports (16) fixed to opposite sides of the base (5b) and extending on the same side of the transmission gearbox to their link (19) to structural supports (20) of the fuselage (6). Elastic links (17) provide the suspension at least in the direction perpendicular to the rods (14). In a variant, the rods (14) are linked to the fuselage (6) being articulated on a transverse flexible bar, equipped as required with a central flapping mass, or on transverse levers articulated onto one another by a central articulation and onto the structure. Application to uni-or bidirectional, focal-point and, as appropriate, antiresonant suspension of helicopter main rotors.