Abstract: A flight test program was conducted to obtain data from an upgraded Gazelle helicopter with an advanced geometry, three bladed rotor. Data were acquired on upper and lower surface chordwise blade pressure, blade bending and torsion moments, and fuselage structural loads. Results are presented from 16 individual flight conditions, including level flights ranging from 10 to 77 m/sec at 50 to 3000 m altitude, turning flights up to 2.0 g, and autorotation. Rotor aerodynamic data include information from 51 pressure transducers distributed chordwise at 75, 88, and 97% radial stations. Individual tranducer pressure coefficients and airfoil section lift and pitching moment coefficients are presented, as are steady state flight condition parameters and time dependence rotor loads. All dynamic data are presented as harmonic analysis coefficients.

Abstract: An aircraft comprising an aerodynamic annulus and a fuselage coupled to the annulus by electromagnetic means to permit rotation of the annulus relative to the fuselage without mechanical engagement upon a "frictionless" bearing. Jets mounted on the annulus rotate the annulus to produce a gyroscopic precession. One or more jets or driven propeller units mounted on the fuselage drive the craft in a selected direction. Direction of the fuselage, and hence the aircraft, may be provided by computer control of the fuselage jets or jets or by means of a rudder. Pivotal jets mounted on the annulus and/or fuselage together with retractable flaps on the annulus provide lift for vertical take-off and hovering.

Abstract: An accurate and efficient numerical solution is developed for predicting high-frequency radiation patterns of antennas mounted on curved surfaces. This solution employs the uniform geometrical theory of diffraction (UTD) and has mainly been used to analyze airborne antenna patterns. In this case the aircraft is modeled in its most basic form so that the solution is applicable to general-type aircraft. The fuselage is modeled as a perfectly conducting composite ellipsoid, whereas, the wings, stabilizers, nose, fuel tanks, and engines, etc. are simulated by perfectly conducting fiat plates. The composite-ellipsoid fuselage model is necessary to simulate successfully the wide variety of real world fuselage shapes. Since the antenna is mounted on the fuselage, it has a dominant effect on the resulting radiation pattern, so it must be simulated accurately, especially near the antenna. Various radiation patterns are calculated for military aircraft, private aircraft, and the space shuttle orbiter. The application of this solution to practical airborne antenna problems illustrates its versatility and design capability. The solution accuracy is verified by the comparisons between calculated and measured data.

Abstract: The flow over a 58-mm-thick uniform-thickness winglike body having a 1.5:1 elliptical leading edge and joined to a large flat plate (representing an aircraft fuselage) is characterized experimentally at freestream velocity 15 m/s, corresponding to Reynolds number 940,000/m, using hot-wire anemometry. The results are presented graphically, and it is found that the horseshoe vortex formed by the separation of the fuselage boundary layer ahead of the wing leading edge is effective in transporting turbulence and modifying the mean-flow characteristics and the turbulent-stress distribution. It is suggested that the slenderness ratio of the leading edge is the dominant factor affecting the strength and location of the vortex.

Abstract: A simplified cylindrical model of an aircraft fuselage is used to investigate the mechanisms of interior noise suppression of the synchrophasing technique. This investigation allows isolation of important parameters to define the characteristics of synchrophasing. The optimum synchrophase angle for maximum noise reduction is found for several interior microphone positions with pure tone source conditions. Noise reductions of up to 30dB are shown for some microphone positions, however, overall reductions are less. A computer algorithm is developed to decompose the modal composition of the cylinder vibration over a wide range of synchrophase angles. The circumferential modal response of the shell vibration is shown to govern the transmission of sound into the cylinder rather than localized transmission.

Abstract: This invention relates to a collapsible fuel storage tank for storing fuel on an aircraft. According to this invention, there is provided an aircraft including a fuselage portion, a wing portion connected thereto and an expandable fuel storage tank disposed adjacent the intersection of the wing portion and the fuselage portion. The fuel storage tank has an external surface which includes spanwise spaced inner and outer edge regions extending in a generally chordwise direction. The inner edge region of the tank extending alongside the fuselage portion and the outer edge region being attached to the wing portion to allow movement of the external surface between an empty position in which it lies generally flush with the surrounding surface of the wing portion and a fuel storing position in which it lies at an acute angle to said surrounding surface forming a fillet region between the wing portion and the fuselage portion. By this arrangement when the external surface is in a fuel storing position the wetted area of the aircraft is decreased, rather than increased. In addition, the external surface moves to its fuel storing position without creating large gaps or slots.

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: In this paper the noise control characteristics of synchrophasing are investigated using a simplified model of an aircraft fuselage. The analysis presented here includes directivity effects of the noise sources and solves in closed form the coupled motion between the interior and exterior acoustic fields and the shell vibrational response. The variation in sound pressure level at various locations inside the shell is studied for various synchrophase angles as well as the shell vibrational response and input power flow in order to uncover the principal mechanisms behind the transmission phenomena.

Abstract: A high speed helicopter, ideally suited for use with rotors which are tip propelled, has an offset flapping hub mounted on a tiltable mast which is located a substantial distance off to the retreading blade side of the fuselage of the helicopter. This causes the fuselage to be located completely under the advancing blades. Individual pitch control motors individually vary the pitch of the rotor blades to control the feathering angle as a function of the angular position so that lift is obtained from the advancing and laterally moving blades only.

Abstract: A wind-tunnel investigation was conducted in which aerodynamic loads were measured on a small-scale helicopter rotor and a body of revolution located close to it as an idealized model of a fuselage. The objective was to study the aerodynamic interactions as a function of forward speed, rotor thrust, and rotor/body position. Results show that body loads, normalized by rotor thrust, are functions of the ratio between free-stream velocity and the hover-induced velocity predicted by momentum theory.

Abstract: An analytical method has been developed for assessing crash dynamics of large transport aircraft, such as in the NASA Controlled Impact Demonstration (CID) jet crash test on December 1, 1984. The DYCAST nonlinear finite-element computer code was used in a series of progressively more difficult tasks to model complete transport aircraft crashes. Single aircraft frames and fuselage section vertical drop tests were modeled and analyzed to obtain comparisons with experimental data and to develop hybrid element crash springs for use in the large CID model. Predictions of crash and acceleration levels from a symmetric CID model agreed well with data from the CID experiment.

Abstract: An explanation is provided for the difference in wake vortex alleviation achieved by roll oscillations in flight tests with B-747 and L-1011 transport aircraft. Both aircraft had their landing flaps extended and several spoilers deployed. Numerical analysis shows that the growth in amplitude of the initial waves in the vortex filament is brought about by the sinusoidal instability. In the case of the B-747, growth is enhanced by a vortex whose strength is about the same as the tip vortex which is shed near the fuselage by the inboard end of the flaps. Conversely, the L-1011 is estimated to shed a negligible fuselage vortex and to have a relatively strong wingtip vortex. These characteristics bring about a rotation and amplification of the initial waves in the vortex filaments in the wake of the B-747 but not in the L-1011 vortex wake. An aircraft following the B-747 would then experience only intermittent encounters with the intense parts of the wake vortices so that the time-averaged wake-induced rolling moment is substantially reduced.

Abstract: A fluid flow control device controllably maintains attached flow in the region of a body having a contour of rapid curvature utilizing tangential fluid discharge slots, positioned just upstream from the separation line, which issue a thin jet sheet to energize the boundary layer and entrain the surrounding flow. When applied to the aft fuselage of an aircraft, the device reduces separation and vortex drag at cruise and provides control forces and moments during low speed operation of the aircraft.

Abstract: A 12-foot long Boeing 707 aft fuselage section with a tapering cross section was drop tested at the NASA Langley Research Center to measure structural, seat, and occupant response to vertical crash laods and to provide data for nonlinear finite element modeling. This was the final test in a series of three different transport fuselage sections tested under identical conditions. The test parameters at impact were: 20 ft/s velocity, and zero pitch, roll, and yaw. In addition, the test was an operational shock test of the data acquisition system used for the Controlled Impact Demonstration (CID) of a remotely piloted Boeing 720 that was crash tested at NASA Ames Dryden Flight Research Facility on December 1, 1984. Post-test measurements of the crush showed that the front of the section (with larger diameter) crushed vertically approximately 14 inches while the rear crushed 18 inches. Analysis of the data traces indicate the maximum peak normal (vertical) accelerations at the bottom of the frames were approximately 109 G at body station 1040 and 64 G at body station 1120. The peak floor acceleration varied from 14 G near the wall to 25 G near the center where high frequency oscillations of the floor were evident. The peak anthropomorphic dummy pelvis normal (vertical) acceleration was 19 G's.

Abstract: A program was performed to develop and demonstrate the impact dynamics and acoustic transmission technology for a composite fuselage which meets the design requirements of a 1990 large transport aircraft without substantial weight and cost penalties. The program developed the analytical methodology for the prediction of acoustic transmission behavior of advanced composite stiffened shell structures. The methodology predicted that the interior noise level in a composite fuselage due to turbulent boundary layer will be less than in a comparable aluminum fuselage. The verification of these analyses will be performed by NASA Langley Research Center using a composite fuselage shell fabricated by filament winding. The program also developed analytical methodology for the prediction of the impact dynamics behavior of lower fuselage structure constructed with composite materials. Development tests were performed to demonstrate that the composite structure designed to the same operating load requirement can have at least the same energy absorption capability as aluminum structure.

Abstract: Recent technological advances in airplane construction techniques and materials employing bonded and milled aluminum skins and composite materials allow for the production of aerodynamic surfaces without significant waviness and roughness, permitting long runs of natural laminar flow (NLF). These advances lead to excellent opportunities for reducing the drag of aircraft by increasing the extent of NLF. The present research effort seeks to refine and validate computational design tools for use in the design of axisymmetric and nonaxisymmetric natural-laminar-flow bodies. The principal tasks of the investigation involve fuselage body shaping using a computational design procedure. Under Phase I SBIR funding for this research, analytical methods were refined and exploratory calculations were conducted to predict laminar boundary- layer behavior on selected body shapes. Using a low-order surface-singularity aerodynamic analysis program pressure distribution, boundary-layer development, transition location and drag coefficient have been obtained for a number of body shapes including a representative business-aircraft fuselage. Extensive runs of laminar flow were predicted in regions of favorable pressure gradient on smooth body surfaces. A computational design procedure was developed to obtain a body shape with minimum drag coefficient having large extent of NLF. Some preliminary results from the design efforts have been obtained and further work is underway. The proposed study has widespread commercial applications. A significant reduction in the drag produced by any airplane can be obtained when extensive runs of natural laminar flow are achieved on its fuselage resulting in improved airplane performance and efficiency.

Abstract: The prediction sequence used in the aircraft noise prediction program (ANOPP) is described. The elements of the sequence are called program modules. The first group of modules analyzes the propeller geometry, the aerodynamics, including both potential and boundary-layer flow, the propeller performance, and the surface loading distribution. This group of modules is based entirely on aerodynamic strip theory. The next group of modules deals with the first group. Predictions of periodic thickness and loading noise are determined with time-domain methods. Broadband noise is predicted by a semiempirical method. Near-field predictions of fuselage surface pressrues include the effects of boundary layer refraction and scattering. Far-field predictions include atmospheric and ground effects.

Abstract: On December 1, 1984, the FAA and NASA conducted a remotely piloted air-to-ground crash test of a Boeing 720 transport aircraft instrumented to measure crash loads of the structure and the anthropomorphic dummy passengers Over 330 time histories of accelerations and loads collected during the Full-Scale Transport Controlled Impact Demonstration (CID) for the 1-sec period after initial impact are presented Although a symmetric 1 deg nose-up attitude with a 17 ft/sec sink rate was planned, the plane was yawed and rolled 13 deg at initial (left-wing) impact The first fuselage impact occurred near the nose wheel well with the nose pitched down 25 deg Peak normal (vertical) floor accelerations were highest in the cockpit and forward cabin near the nose wheel well and were approximately 14G The remaining cabin floor received normal acceleration peaks of 7G or less The peak longitudinal floor accelerations showed a similar distribution, with the highest (7G) in the cockpit and forward cabin, decreasing to 4G or less toward the rear Peak transverse floor accelerations ranged from about 5G in the cockpit to 1G in the aft fuselage

Abstract: Tilt-rotor flutter control under cruising operation is analyzed. The rotor model consists of a straight fixed wing, a pylon attached to the wingtip, and a three-blade rotor. The wing is cantilevered to the fuselage and is allowed to bend forward and upward. It also has a torsional degree of freedom about the elastic axis. Each rotor blade has two bending degrees of freedom. Feedback of wingtip velocity and acceleration to cyclic pitch is investigated for flutter control, using strip theory and linearized equations of motion. To determine the feedback gain, an eigenvalue analysis is performed. A second, independent, timewise calculation is conducted to evaluate the control law while employing more sophisticated aerodynamics. The effectiveness of flutter control by cyclic pitch change was confirmed.

Abstract: The controlled impact demonstration (CID) test of a transport aircraft took place on December 1, 1984, crashing at a prepared site on Rogers Dry Lakebed, Edwards Air Force Base, California. The demonstration was a setback for the antimisting kerosene (AMK) researchers. The impact conditions, considerably different from the planned scenario, exposed large quantities of degraded AMK and hydraulic fluid and caused unexpectedly hot ignition sources, bulk loss of fuel from the right wing, airflow patterns over the wings and fuselage that were untested on AMK, and fuel intrusion into the lower fuselage. The test was much more severe than planned and is generally considered to be unrepresentative of the type of survivable crash that would benefit from AMK. Ninety-seven percent of the sensors on the fuselage and wing structure, seats, dummies, restraint systems, galley, and bins were active at impact. A wealth of sensor data was collected from this once-in-a-lifetime research test. The flight data recorder experiments on board were also generally successful.

Abstract: An aircraft cargo container having sides, inboard and outboard ends, a horizontal top and a horizontal bottom. The bottom is rectangular and provided with casters located in corner recesses. The inboard end and both sides of the container are substantially vertical, while the outboard end substantially conforms to the curvature of the aircraft fuselage cabin cross section. The inboard and outboard ends are so sized that the container will freely pass through a standard left side passenger entry door. The sides are so dimensioned that when two containers are located end-to-end with their inboard ends opposed, they will substantially fill the aircraft fuselage cabin cross section with clearance between themselves and between themselves and the aircraft fuselage, so that a plurality of containers can be arranged within the aircraft in two longitudinal rows, the containers of each row having adjacent sides opposed. Each container has a door in one of its sides. The container bottom provides flanges along the container ends cooperating with side guide rails and a center guide rail assembly mounted in the aircraft. The container bottom also provides flanges along the container sides, engageable by fore and aft restraints.

Abstract: In the case of a control device for collective and cyclic adjustment of rotor blades having a swash plate (3) which is coaxial with respect to the rotor drive shaft (2), this swash plate (3) is supported (either) on the rotor side (or on the fuselage side) by a universal joint (8) consisting of an assembly of four hinged bearings (9, 10) which can move in an angle in all directions and are offset through 90 DEG with respect to one another around the circumference of the swash plate in a plane which is parallel to the rotor rotation plane, that pair of mutually opposite hinged bearings (9) which provides support on the rotor side (or fuselage side) being arranged such that it can move in the direction in which the rotor drive shaft (2) extends.

Abstract: A series of low-speed wind tunnel tests on a generic airplane model with a cylindrical fuselage were made to investigate the effects of forebody shape and fitness ratio, and fuselage/wing proximity on static and dynamic lateral/directional stability. In addition, some preliminary testing to determine the effectiveness of deflectable forebody strakes for high angle of attack yaw control was conducted. During the stability investigation, 11 forebodies were tested including three different cross-sectional shapes with fineness ratios of 2, 3, and 4. In addition, the wing was tested at two longitudinal positions to provide a substantial variation in forebody/wing proximity. Conventional force tests were conducted to determine static stability characteristics, and single-degree-of-freedom free-to-roll tests were conducted to study the wing rock characteristics of the model with the various forebodies. Flow visualization data were obtained to aid in the analysis of the complex flow phenomena involved. The results show that the forebody cross-sectional shape and fineness ratio and forebody/wing proximity can strongly affect both static and dynamic (roll) stability at high angles of attack. These characteristics result from the impact of these factors on forebody vortex development, the behavior of the vortices in sideslip, and their interaction with the wing flow field. Preliminary results from the deflectable strake investigation indicated that forebody flow control using this concept can provide very large yaw control moments at stall and post-stall angles of attack.

Abstract: A device for tow vehicles intended for maneuvering commercial aircraft, where the type of aircraft to be towed is identified and the turn-out angle between the longitudinal axes of the vehicle and the aircraft is determined and monitored during the towing process. The type of aircraft may be ascertained by measuring the width of the nose gear assembly. The turn-out angle can be calculated by measuring the relative distances between opposing sensors on the tow vehicle and the fuselage utilizing a characteristic hull shape of the particular type of aircraft.

Abstract: Unsteady transonic flow calculations are presented for wing/fuselage configurations. Calculations are performed by extending the XTRAN3S (version 1.5) unsteady transonic small-disturbance code to allow the treatment of a fuselage. The research was conducted as part of a larger effort directed toward developing the capability to treat a complete flight vehicle. Details of the XTRAN3S fuselage modeling are discussed in the context of the small-disturbance equation. Transonic calculations are presented for two wing/fuselage configurations with leading-edge sweep angles of 0 and 36.65 deg, the results of which compare well with available experimental steady-pressure data. Unsteady calculations are performed for simple bending and torsion modal oscillations of the wing. Comparisons of sectional lift and moment coefficients for the wing-alone and wing/fuselage cases reveal effects of fuselage aerodynamic interference on the unsteady wing loading. Tabulated generalized aerodynamic forces, typically used in flutter analyses, indicate small changes in the real (in-phase) component and as much as a 30% change in the imaginary (out-of-phase) component when the fuselage is included in the calculation. These changes result in a 2-5% increase in total magnitude and a several degree increase in phase.

Abstract: This invention is an aircraft 10 with a system for increasing the lift-drag ratio over a broad range of operating conditions. The system positions the engines and nacelles 15 over the wing 12 in such a position that gains in propeller 16 efficiency is achieved simultaneously with increases in wing lift and a reduction in wing drag. Adverse structural and torsional effects on the wings 12 are avoided by fuselage mounted pylons which attach to the upper portion of the fuselage 11 aft of the wings. Similarly, pylon-wing interference is eliminated by moving the pylons to the fuselage. Further gains are achieved by locating the pylon surface area aft of the aircraft center-of-gravity, thereby augmenting both directional and longitudinal stability. This augmentation has the further effect of reducing the size, weight and drag of empennage components 13. The combination of design changes results in improved cruise performance and increased climb performance while reducing fuel consumption and drag and weight penalties.