Abstract: A computational method has been developed to treat the unsteady aerodynamic interaction between a helicopter rotor, wake, and fuselage. Two existing codes, a lifting line-prescribed wake rotor analysis and a source panel fuselage analysis, were modified and coupled to allow prediction of unsteady fuselage pressures and airloads. A prescribed displacement technique was developed to position the rotor wake about the fuselage. Also coupled into the method were optional blade dynamics or rigid blade performance analyses to set the rotor operating conditions. Sensitivity studies were performed to determine the influence of the wake and fuselage geometry on the computational results. Solutions were computed for an ellipsoidal fuselage and a four bladed rotor at several advance ratios, using both the classical helix and the generalized distorted wake model. Results are presented that describe the induced velocities, pressures, and airloads on the fuselage and the induced velocities and bound circulation at the rotor. The ability to treat arbitrary geometries was demonstrated using a simulated helicopter fuselage. Initial computations were made to simulate the geometry of an experimental rotor-fuselage interaction study performed at the Georgia Institute of Technology.

Abstract: An aeroelastic analysis method for fighter aircraft operating at extreme flight conditions has been developed and tested. The method involves the use of state-of-the-art zonal grid generation methods, three-dimensional Reynoldsaveraged Navier-Stokes analysis, and linear structures to analyze the flow over complex, flexible aircraft. The main objective of this effort is to develop a method capable of analyzing aircraft operating at flight conditions where vortices, strong shock waves, separated flow, and even highly unsteady flow may be present. The present application focuses on the static aeroelastic analysis of fighter aircraft operating at high angle of attack and high transonic Mach number. The developed method has been compared against static aeroelastic wind-tunnel data on an aeroelastically tailored wing/fuselage configuration, and the results are very encouraging.

Abstract: A high-capacity aircraft fuselage having two lobes placed side by side and assembled together along two longitudinal lines of junction is designed to facilitate passenger movement from one lobe to the other in accordance with aviation safety regulations. The fuselage includes a top longeron (28) for reinforcing the top line of junction, a floor (23) providing a separation between an upper internal space which can serve as a cabin (24) and a lower internal space (25) which can serve as a hold, a girder (32) for supporting the cabin floor and connecting it to the bottom portion of the fuselage, rows of seats extending from one side of the bilobed body to the other, a bottom longeron (29) for reinforcement along the bottom line of intersection, and posts (31) disposed in spaced relation in the rows of seats for connecting the top longeron (28) to structural members located in the lower region of the fuselage (32, 29).

Abstract: : Investigations of tail buffet on the CF-18 have been conducted at the National Aeronautical Establishment (NAE) and the Aerospace Engineering Test Establishment (AETE). Flow visualization of the vortex burst phenomenon was carried out in a low speed water tunnel using a modified 1/72 scaled plastic model. In wind tunnel tests, a rigid 6% model was used for measurements in the NAE 5ft x 5ft Trisonic Tunnel. Unsteady pressure measurements on the vertical fin were made by means of 24 fast response transducers on each surface. Results of the acceleration experienced by the fin are presented Vortex flow structure was studied with the aid of a 49 pressure-sensor-rake mounted behind the fin. In addition to measuring steady pitot pressure values to deduce pressure contours, unsteady pressure fluctuations were obtained from 13 fast response transducers. The LEX was also instrumented with pressure orifices and fast response transducers. The investigation was carried out with LEX fences 'on' and 'off' to note their effect on tail buffet loads. Flight tests have been conducted at AETE on a test aircraft with accelerometers installed on the vertical fins and horizontal stabilators and strain gauges mounted on the aft fuselage structures and fin root attachment stubs. Flight test data are presented showing the effectiveness of the LEX fence in reducing aft fuselage structural to buffet loads.

Abstract: An automated system is disclosed for rapidly inspecting a complete aircraft and detecting cracks in the airframe. For inspecting the aircraft fuselage, the system incorporates a framework disposed over the fuselage and extending along a major portion of the length of the fuselage. Beams having suction devices and acoustic sensors are movably attached to the framework and are moved into a position adjacent to the outer surface of the fuselage when the aircraft has been located within the framework. The suction devices attach the beams to the fuselage surface along fuselage panel joints or other areas to be inspected. Several acoustic sensors, located on each beam, are connected to a device for analyzing and recording or visually displaying the signals generated by the sensors upon the detection of noise generated by the formation or propagation of cracks. In order to simulate the loads on the fuselage encountered during flight, the interior of the fuselage is pressurized via the aircraft engines or an external pressurization source. The system according to the invention may also be used to inspect the aircraft wings by placing inflatable bags beneath the wing and inflating them so as to exert upward loads on the wing. Additional inflatable bags are placed between the fuselage and a framework extending over the aircraft fuselage such that, when inflated, they exert downward loads on the upper surface of the aircraft fuselage. The fuselage is pressurized to prevent collapse. Mounting beams having acoustic sensors are attached via suction devices to the aircraft wing at joints, or other areas to be inspected.

Abstract: The injection of thin, high-momentum jets of air into the fuselage forebody boundary layers of the F-18 aircraft is explored numerically as a means of controlling the onset of fuselage vortices and of generating yaw control forces. The study was carried out for an angle of attack of 30 deg with symmetrical and asymmetrical blowing configurations. One-sided blowing results in a strongly asymmetrical flow pattern in the fore portion of the fuselage, leading to a net lateral force.

Abstract: An all-wing aircraft is disclosed that has novel foreplane and depending aftplane, a center wing section and outer wing panel flying surfaces which cooperate aerodynamically to eliminate the need for conventional fuselage and tail structures. The foreplanes are strategically located to create a positive pitching moment which is sufficient to significantly reduce elevator forces and to balance the negative pitching moment induced by the outer wing panels and the downwardly extending aftplane units that provide static and dynamic pitch and yaw stability. At the same time, the foreplane structures serve as an unobstructed means to mount engines forward on the airframe to established a forward empty center of gravity. Additionally the aftplane structure serves as a means to mount main landing gears, elevators, ruddervators and to provide structural means for interconnecting outboard wing sections of the airplane. The all-wing design is of a nature permitting adoption of the principles thereof in either a multi-engine airplane or single-engine airplane. In both instances, a center wing section, foreplane and aftplane structures and outer wing sections cooperate to provide stability in all three axes of movement of the airplane while decreasing fuel burn by virtue of an improved lift to drag ratio and empty weight reduction by removal of the conventional fuselage and tail structures.

Abstract: An active noise control model has been evaluated for reducing aircraft interior noise. The structural noise transmission properties of an aircraft fuselage were modelled as a flexible cylinder excited by external acoustic dipoles simulating the noise produced by twin propellers. The amplitudes of an internal distribution of monopole control sources were determined such that the area-weighted mean square acoustic pressure was minimized in the propeller plane. The noise control model was evaluated at low frequencies corresponding to the blade passage frequency and first few harmonics of a typical turbo-prop aircraft. Interior noise reductions of 20 25 dB were achieved, over a substantial region of the cylindrical cross-section, with just a few monopole control sources. The most favorable interior noise reductions were achieved when the active noise control model was used in combination with propeller source phasing.

Abstract: Transonic Navier-Stokes flow simulations are presented for the F-16A fighter aircraft using a zonal grid approach. This approach subdivides the physical space about the aircraft into an ensemble of simple geometric shapes, thus mitigating many of the difficulties of generating a single grid about a complex shape, e.g., providing adequate grid refinement near all body surfaces to capture the boundary layer. Information is propagated between zones via grid overlapping and a spatial interpolation procedure. Computational Cp compare well with experimental values on the wing, horizontal and vertical tails, fuselage centerline, and the inlet/diverter region. The average y+ one grid point off the wing is 3. The experimental lift is underpredicted by 2.6%, and the experimental drag is overpredicted by 1.6%. The flexibility of the zonal approach is demonstrated by adding additional zones inside the inlet up to the compressor face to model flow spillage, and downwind of the exhaust nozzle to model power-on conditions. Computations are also presented for the F-16A in sideslip. These results demonstrate that the present zonal approach provides a flexible and viable means of simulating flowfields about complex geometries.

Abstract: This invention relates to a lift device adapted to transfer wheelchair passengers to commuter aircraft. A base and optional four post system provide support for a vertically moveable platform. An apron, which can be used either as a ramp from lowered platform to ground or a bridge from elevated platform to aircraft is provided. The platform is preferably raised and lowered using ball or acme screw-bushing brackets. The platform can be used to load and unload passengers from rear and front door commuter aircraft, on either side of the fuselage. The lift can be incorporated into a wheeled chassis, which can be lowered into ground contact at the position of use. Optionally extendible stairs are present to allow other passengers to enter or leave the aircraft using said platform and apron.

Abstract: In support of the NASA High Alpha Technology Program, Navier-Stokes solutions have been obtained using the Chimera overset grid scheme for flow over the wing, fuselage, and wing leading-edge extension (LEX) of the F/A-18 High Alpha Research Vehicle (HARV) at high incidence. Solutions are also presented for flow over the fuselage forebody at high angles of attack. The solutions are for turbulent flows at high-Reynolds-number flight-test conditions, and are compared with available qualitative and quantitative experimental data. Comparisons of predicted surface flow patterns, off-surface flow visualization, and surface-pressure distributions are in good agreement with flight-test data. The ability of the numerical method to predict the bursting of the LEX vortex as it encounters the adverse pressure gradient field of the wing is demonstrated, and the capability of predicting high-angle-of-attack aerodynamics around realistic aircraft configurations is established.

Abstract: A heli-hover amphibious craft which includes a main hover craft fuselage body structure including a single hull provided with a deck. The hull is defined by a compartmental under base cavity extending downward with divisional and perimeter walls formed by air cushion containing structure for containing suspension air. The fuselage body is provided with at least one built-in duct extending through said deck to said base cavity. At least one deck mounted fan is provided for delivering pressurized air through the duct. A superstructure is attached to the fuselage body. At least one horizontally rotating heli-rotor assembly is supported well above the deck by the superstructure, with the heli-rotor assembly being driven by a substantially vertically fixed shaft. An anti-torque device provided on an aft portion of the fuselage body. A drive mechanism is operatively connected to the heli-rotor assembly and the anti-torque device for propelling the craft.

Abstract: A mathematical model of a helicopter system with a single main rotor that includes rigid, hinge-restrained rotor blades with flap, lag, and torsion degrees of freedom is described. The model allows several hinge sequences and two offsets in the hinges. Quasi-steady Greenberg theory is used to calculate the blade-section aerodynamic forces, and inflow effects are accounted for by using three-state nonlinear dynamic inflow model. The motion of the rigid fuselage is defined by six degrees of freedom, and an optional rotor rpm degree of freedom is available. Empennage surfaces and the tail rotor are modeled, and the effect of main-rotor downwash on these elements is included. Model trim linearization, and time-integration operations are described and can be applied to a subset of the model in the rotating or nonrotating coordinate frame. A preliminary validation of the model is made by comparing its results with those of other analytical and experimental studies. This publication presents the results of research compiled in November 1989.

Abstract: In the case of a pressure wall for a fuselage of an aircraft, for pressure-tight axial termination of a fuselage area which can be internally pressurised, the forces resulting from the internal pressure being introduced into the fuselage structure, the invention comprises the provision of a network 5, which is formed from tensile elements 8, 9, to absorb the compression forces, and a membrane 4 in order to ensure pressure tightness. In this case, it is particularly advantageous that, in addition to the advantages achieved as a result of the object, the invention results in considerably improved accessibility to load-bearing parts for inspection and repair work, possibly even in flight.

Abstract: A tail sitter aircraft of a type that takes off and lands on its tail section with the fuselage pointed vertically upward has landing gears located on its rearward end. Each of the landing gears is mounted on a support axis which is offset from the fuselage axis and which intersects a wheel axle at a 90 degree angle. The landing gears are located on the tail section and spaced around the fuselage axis. Two of the landing gears which are opposite each other will be locked so that they can roll only along a single straight line. The other two landing gears, which are also spaced opposite each other, are locked so that they can roll only on a single straight line. The straight lines are perpendicular to each other.

Abstract: A tail sitter airplane will take off and land on its tail section, with its fuselage and nose pointed up. The airplane has a single driven propeller mounted to the nose section. Wings extend outward from the fuselage. Four airfoils locate in the tail section, two horizontal and two vertical. Each has a control surface. Four additional airfoils locate in a forward section, behind the propeller and in front of the wings. Two of the airfoils in the forward section are horizontal and two vertical. Movable control surfaces on these airfoils control the flight during takeoff and landing. The airfoils have chord lengths selected to remove the twist from the slip-stream from the propeller, thereby balancing the torque from the propeller.

Abstract: General Electric has developed an Ultra-high bypass open rotor engine(UDF@’ engine)When Boeing was studying the use this engine with counter rotating propulsors, possible noise increase due to installation on the aft end of the fuselage was one of the concerns. Such noise increase could be due to the inflow distortion and acoustic diffraction. To determine the installation effects, a test of a scale model of the U D e r o puisor was conducted at the German-Dutch wind tunnel (DNW) in Holland. To simulate installation effects, models of pylon, fore-shortened-aft-fuselage and empennage were used. These could be added in various combinations to enable effects of individual items to be determined. A pylon trailing edge blowing technique was tested to reduce the impact of installation effects. The test results showed that the addition of the pylon upstream of the propulsor could increase the front rotor blade passage frequency(BPF) by as much as 10-12 dB with negligible effect on interaction tones. Addition of the fuselage increased the front rotor BPF by an additional 2-3 dB and also impacted the interaction tones. The root-flow at the attachment point of the pylon to the fuselage was also seen as a source of noise increase. BPF of the front rotor was seen to increase monotonically as the pylon was brought close to the propulsor and when the fuselage was moved closer to the propulsor tip. Pylon trailing edge blowing technique was found to be successful in reducing the pylon wake and the root-flow distortion and was capable of bringing the installed configuration front BPF levels back down close to the isolated propulsor levels. Comparison of data taken at Boeing Transonic Wind Tunnel facility with that taken at DNW agreed reasonably well providing confidence in repeatability of model propulsor data and in noise data taken at BTWT with acoustically treated tunnel walls.

Abstract: An aircraft having an upswept tail section fuselage includes a single pair of large vortex generators mounted in the vicinity of the break in the fuselage ahead of or at the beginning of the tail upsweep, each vortex generator being mounted on a side of and adjacent the bottom of the fuselage. The vortex generators, which may be plates or fins, develop strong transverse outflow from the vertical plane of symmetry that relieves or delays the tendency to flow separation by acting on the external flow field while at the same time energizing the boundary layer to increase its resistance to separation. The vortex generator may be thin or slightly thickened, flat or airfoil shaped, and may have triangular, straight, tapered, or reverse tapered planforms and may be rigid or flexible. Either one or both edges may be blunted, although sharp edges are preferred.

Abstract: An aircraft including a fuselage and a double wing unit which is pivotably mounted onto the fuselage and arranged to selectably assume a first orientation, generally parallel to the fuselage, during storage and transport and a second orientation, generally perpendicular to the fuselage, for flight.

Abstract: The effects of a fuselage and its boundary layer on sound propagation to the fuselage surface and on sound scattering in the farfield were analyzed. A hard-wall infinite cylinder with a boundary layer of both velocity and temperature variations was modeled to simulate the fuselage of an aircraft in flight. Examples for a monopole noise source outside the boundary layer showed considerable noise attenuation on the cylindrical surface forward of the source and much less effect on the downstream side. Data from a transonic wind tunnel test showed the same trends. For enroute and airport community noise, the boundary layer alters the interference pattern caused by the fuselage.

Abstract: A supersonic guided missile has a fuselage terminating at the front in a nose and at the rear in a base and is provided externally with fixed rear planes. At a longitudinal distance from the center of gravity is at least one spoiler mobile transversely between a configuration retracted inside the fuselage and an active configuration in which the spoiler projects laterally from the fuselage.

Abstract: Active noise control has been shown to be a promising solution for reducing noise levels in turboprop aircraft. The first part of this paper shows the performance of a broadband active noise control system based upon an off-line frequency domain model obtained from response measurements. A laboratory test in an 18 passenger airplane fuselage resulted in 10-20 dB reduction of broadband deterministic noise over a substantial frequency range

Abstract: The velocity field at locations above the plane of the rotor disk of a generic helicopter configuration has been measured experimentally and modeled using two computational models. The first computational model represents only the portion of the velocity field due to the fuselage, using only the potential due to freestream onset flow. The second model represents all elements of the problem in an incompressible field: the rotor, the rotor wake, and the non-lifting fuselage. The first model is used to assess the fidelity requirement for fuselage paneling. Comparisons are made between a much simplified body shape, an accurate body shape, a body shape with a crude hub/shaft model and wind-tunnel measurements. The second model is used to assess the relative magnitude of the velocity perturbations at the rotor disk due to the effects of the fuselage. A fully interactive time-stepping wake method is used to predict the velocities at the rotor disk. These velocities are computed with and without the fuselage in the model and compared with measured wind-tunnel inflow velocities. The difference between the velocities computed with and without the fuselage is then compared with the perturbation velocity field computed by the first method.

Abstract: A procedure for designing a fuselage having a prescribed effective area distribution computed from -90 deg Mach slices is described. This type of calculation is an essential tool in designing a complete configuration with an effective area distribution that corresponds to a desired sonic boom signature shape. Sample calculations are given for M=2 and M=3 designs. The examples include fuselages constrained to have circular cross sections and fuselages having cross sections of arbitrary shape. It is found that, for a prescribed effective area distribution having sharp variations, the iterative procedure converges to a smoothed approximation to the prescribed distribution. For a smooth prescribed area distribution, the solution is not unique.

Abstract: A vertical take-off and landing aircraft is configured so as to provide, when landing tail-first with its fuselage (1) in a generally vertical attitude, a touchdown area (21) at the tail of the aircraft at a position offset from a line extending along the length of the fuselage through the centre of gravity (23) of the aircraft such that after the touchdown area contacts a landing surface the aircraft topples under the action of gravity to bring an undercarriage (11, 13) of the aircraft into contact with the landing surface, thereby to attain a stable landed position.