Abstract: A vertical take-off and landing vehicle comprised of a fuselage having a front, a rear, and two lateral sides and a set of four thrusters set to the front, the left, the right, and the rear of said fuselage. The thrusters are either independently powered thrusters or could utilize a single power source. The thrusters, which are ducted fan units capable of providing a vertically upward force to the aircraft, are provided with such redundancy that the aircraft can hover with up to two thrusters inoperative. The thrusters are comprised of a set of two counter rotating propellers both of which creates lift. The two counter rotating propellers cancel out the torque effect normally created by using only one propeller. The Ducted fan units being movable between a first position in which they provide vertical lift and a second position in which they provide horizontal thrust using a set of servos and gears.

Abstract: *† This paper introduces an approach to effectively model the nonlinear aeroelastic behavior of fully flexible aircraft. The study is conducted based on a nonlinear strainedbased finite element framework in which the developed low-order formulation captures the nonlinear (large) deflection behavior of the wings, and the unsteady subsonic aerodynamic forces acting on them. Instead of merely considering the nonlinearity of the wings, the paper will allow all members of the vehicle to be flexible. Due to their characteristics of being long and slender structures, the wings, tail, and fuselage of highly flexible aircraft can be modeled as beams undergoing three dimensional displacements and rotations. The cross-sectional stiffness and inertia properties of the beams are calculated along the span, and then incorporated into the 1-D nonlinear beam model. Finite-state unsteady subsonic aerodynamic loads are incorporated to be coupled with all lifting surfaces, so as to complete the state space aeroelastic model. Different Sensorcraft concepts are modeled and studied, including conventional single-wing and joined-wing aircraft configurations with flexible fuselage and tail. Based on the proposed models, roll responses and stabilities are studied and compared with linearized and rigidized models. At last, effects of the flexibility of the fuselage and tail on the roll maneuver and stability of the aircraft are presented.

Abstract: This paper presents the development and validation of a CFD method suitable for analysing the flow around a complete helicopter. Emphasis is placed on the detailed validation for a range of flow cases designed to assess the predictive capability of the method. Several fundamental flow cases are presented including tip flows, blade-vortex interaction and threedimensional dynamic stall. Solutions for isolated rotors are also obtained both at hover and forward flight conditions. The detailed validation of the method for fundamental flows was found to be beneficial when more complex cases were considered like the flow around a helicopter fuselage or the rotor/fuselage interaction. Throughout, experimental data available in the public domain has been used and a substantial validation database has been compiled. Overall, the method was found capable of capturing the basic flow characteristics, however, the need for detailed experimental data has been identified as a number one priority in order to assess the method in predicting the interactions between the main rotor the fuselage and the tail rotor at a wide range of conditions. Introduction In contrast to CFD solutions for complete fixed-wing aircraft, which appear frequently in the literature, the case of a full helicopter still remains rare [1,2,3]. There are several reasons for this, the main being the complexity of the flow around a helicopter which can challenge modern CFD methods even when the conditions considered are well within the design envelope. A summary of the flow phenomena associated with helicopter flows can be found in the review paper by Conlisk [3] and is discussed in detail in specialised rotorcraft textbooks [4,5,6,7]. For this work three fundamental flow phenomena have been identified as key for CFD validation. These include the formation of the tip vortex [8-9], the threedimensional dynamic stall [10-11] and the blade-vortex interaction [12-13]. In addition, the CFD analysis of the flow around isolated rotors in hover and forward flight conditions requires specific formulation of the governing equations to account for the rotation and the actuation of the blades which is not the case for the fixed-wing problem. For this work several flow cases have been considered based on public-domain experimental data [14-18]. 1 Senior Lecturer, gbarakos@aero.gla.ac.uk 2 Post-Doctoral Research Associate 3 Reader 4 Principal Engineer, PhD Candidate

Abstract: : The flight path dynamics of aircraft are often characterized by the presence of a right-half plane zero in the elevator-to-flight path angle transfer function. For most aircraft, the frequency of this zero is high enough that it does not limit the bandwidth of the flight control system. This is not the case, however, with air-breathing hypersonic aircraft. This class of aircraft is characterized by unstable longitudinal dynamics, strong loop interactions, and the presence of non-minimum phase transmission zeros. In the case of flight-path angle and velocity control, the presence of a low frequency transmission zero severely limits the achievable bandwidth. We show that the frequency of the zero is related to the instantaneous center-of-rotation of the aircraft, which is dependent upon the amount of lift produced by the longitudinal control effectors. In order to improve flight-path control, we investigate the feasibility of an aircraft configured with redundant pitch control effectors. The additional effector moves the instantaneous center-of-rotation, and as a result, the location of the zero. The trade-off is that the path-attitude decoupling inherent in hypersonic aircraft becomes more pronounced. Results are given for both a rigid hypersonic aircraft model and a model that includes the effects of the first fuselage bending mode.

Abstract: An aircraft having a vertical take-off and landing (VTOL) propulsion system. The aircraft includes a fuselage, the VTOL propulsion system, at least one forward thruster, a power source used for both the VTOL propulsion system and forward thruster, fore and aft wings and a plurality of spars attached to and spanning the space between the two wings. The VTOL propulsion system includes a plurality of VTOL cells (including a motor, motor controller, and propeller) attached in a spaced relation along each spar. The VTOL cells are used exclusively for vertical flight or hovering and are powered down as the aircraft develops forward flight velocity and corresponding wing lift. During forward flight the VTOL propellers are articulated to allow the aircraft to take on a low drag configuration. The present invention is suitable for use in manned or un-manned aircraft of any scale.

Abstract: The demanding task of landing a maritime helicopter on a ship at sea is constrained by an operational envelope that places limits on wind direction and speed. An operational envelope is normally developed by first-of-class flight trials, but flight testing is an expensive means of qualifying a helicopter for shipborne operations. This paper describes the development of an experimentally-based simulation method which is intended to complement flight testing and mitigate its cost. The basis of the methodology lies with establishing correlations of unsteady aerodynamic fuselage loads (measured in a wind tunnel) with pilot workload (obtained by flight test), assessed for cases where airwake turbulence is chiefly responsible for the workload. With these correlations, contours of unsteady fuselage loading can assist with the definition of the operational envelope. An experiment was conducted to measure the unsteady aerodynamic fuselage loads in a wind tunnel. In this investigation the fuselage of a Sea King helicopter was immersed in both the downwash of a spinning main rotor and the airwake of a Canadian Patrol Frigate. Measurements of unsteady side force, yawing moment and drag force were made over a combination of wind directions, speeds, and hover positions. The results indicate that a spinning main rotor generating appropriate levels of thrust is a necessary feature of the wind-tunnel simulation. Specifically, in comparison to the rotorless case unsteady loading at low hover over the flight deck was found to increase to the levels of unsteadiness that exist at high hover, and the variation of unsteady loading with wind speed is changed by the interaction of the ship airwake and rotor downwash. Generally the unsteady loading increases with the additional influence of main rotor downwash compared to the baseline, rotorless case. The wind tunnel data, particularly side Assistant Research Officer Senior Research Officer Presented at the 29th European Rotorcraft Forum, Friedrichshafen, Germany, September 16-18, 2003. Copyright c 2003 by the National Research Council Canada. force and drag, are then shown to correlate well with flight-test derived operational limits. The correlation of unsteady yawing moment with the operational limit is less straight forward.

Abstract: This paper demonstrates the extension of error estimation and adaptation methods to parallel computations enabling larger, more realistic aerospace applications and the quantification of discretization errors for complex 3-D solutions. Results were shown for an inviscid sonic-boom prediction about a double-cone configuration and a wing/body segmented leading edge (SLE) configuration where the output function of the adjoint was pressure integrated over a part of the cylinder in the near field. After multiple cycles of error estimation and surface/field adaptation, a significant improvement in the inviscid solution for the sonic boom signature of the double cone was observed. Although the double-cone adaptation was initiated from a very coarse mesh, the near-field pressure signature from the final adapted mesh compared very well with the wind-tunnel data which illustrates that the adjoint-based error estimation and adaptation process requires no a priori refinement of the mesh. Similarly, the near-field pressure signature for the SLE wing/body sonic boom configuration showed a significant improvement from the initial coarse mesh to the final adapted mesh in comparison with the wind tunnel results. Error estimation and field adaptation results were also presented for the viscous transonic drag prediction of the DLR-F6 wing/body configuration, and results were compared to a series of globally refined meshes. Two of these globally refined meshes were used as a starting point for the error estimation and field-adaptation process where the output function for the adjoint was the total drag. The field-adapted results showed an improvement in the prediction of the drag in comparison with the finest globally refined mesh and a reduction in the estimate of the remaining drag error. The adjoint-based adaptation parameter showed a need for increased resolution in the surface of the wing/body as well as a need for wake resolution downstream of the fuselage and wing trailing edge in order to achieve the requested drag tolerance. Although further adaptation was required to meet the requested tolerance, no further cycles were computed in order to avoid large discrepancies between the surface mesh spacing and the refined field spacing.

Abstract: Marketing surveys of traditional business aircraft operators have shown that there is significant interest customer interest in the development of a supersonic business jet. One requirement for such an aircraft is that it be capable of unrestricted supersonic flight over land. This requirement has focused considerable attention on sonic boom minimization for a class of 100,000 pound, Mach 1.6 to 2.0 cruise airplanes. Configuring the design with sufficient low boom characteristics while at the same time including enough cabin volume has proven difficult. Gulfstream has developed a Quiet Spike concept that significantly improves the sonic boom signature while at the same time allowing for additional fuselage volume. The multi stage extendable nose spike projects out of, and retracts into, the forward fuselage of the airplane. Wind tunnel tests of supersonic equivalent area distributions have validated the near field aerodynamics of the spike concept. CFD analysis of a wing/body configured with a multi stage spike have extended the analysis beyond idealized equivalent area distributions and shown a great potential for tailoring the shape of the sonic boom ground signature while simultaneously allowing greater flexibility in the fuselage design.

Abstract: The present paper addresses the problem of spin recovery of an aircraft as a nonlinear inverse dynamics problem of determining the control inputs that need to be applied to transfer the aircraft from a spin state to a level trim flight condition. A stable, oscillatory, flat, left spin state is first identified from a standard bifurcation analysis of the aircraft model considered, and this is chosen as the starting point for all recovery attempts. Three different symmetric, level-flight trim states, representative of high, moderate, and low-angle-of-attack trims for the chosen aircraft model, are computed by using an extended-bifurcation-analysis procedure. A standard form of the nonlinear dynamic inversion algorithm is implemented to recover the aircraft from the oscillatory spin state to each of the selected level trims. The required control inputs in each case, obtained by solving the inverse problem, are compared against each other and with the standard recovery procedure for a modern, low-aspect-ratio, fuselage heavy configuration. The spin recovery procedure is seen to be restricted because of limitations in control surface deflections and rates and because of loss of control effectiveness at high angles of attack. In particular, these restrictions adversely affect attempts at recovery directly from high-angle-of-attack oscillatory spins to low-angleof-attack trims using only aerodynamic controls. Further, two different control strategies are examined in an effort to overcome difficulties in spin recovery because of these restrictions. The first strategy uses an indirect, two-step recovery procedure in which the airplane is first recovered to a high- or moderate-angle-of-attack level-flight trim condition, followed by a second step where the airplane is then transitioned to the desired low-angle-of-attack trim. The second strategy involves the use of thrust-vectoring controls in addition to the standard aerodynamic control surfaces to directly recover the aircraft from high-angle-of-attack oscillatory spin to a low-angle-of-attack level-flight trim state. Our studies reveal that both strategies are successful, highlighting the importance of effective thrust management in conjunction with suitable use of all of the aerodynamic control surfaces for spin recovery strategies.

Abstract: Structural analysis and design of efficient pressurized fuselage configurations for the advanced Blended-Wing-Body (BWB) flight vehicle is a challenging problem. Unlike a conventional cylindrical pressurized fuselage, stress level in a box type BWB fuselage is an order of magnitude higher, because internal pressure primarily results in bending stress instead of skin-membrane stress. In addition, resulting deformation of aerodynamic surface could significantly affect performance advantages provided by lifting body. The pressurized composite conformal multi-lobe tanks of X-33 type space vehicle also suffered from similar problem. In the earlier BWB design studies, Vaulted Ribbed Shell (VLRS), Flat Ribbed Shell (FRS); Vaulted shell Honeycomb Core (VLHC) and Flat sandwich shell Honeycomb Core (FLHC) concepts were studied. The flat and vaulted ribbed shell concepts were found most efficient. In a recent study, a set of composite sandwich panel and cross-ribbed panel were analyzed. Optimal values of rib and skin thickness, rib spacing, and panel depth were obtained for minimal weight under stress and buckling constraints. In addition, a set of efficient multi-bubble fuselage (MBF) configuration concept was developed. The special geometric configuration of this concept allows for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls, while the outer-ribbed shell prevents buckling due to external resultant compressive loads. The initial results from these approximate finite element analyses indicate progressively lower maximum stresses and deflections compared to the earlier study. However, a relative comparison of the FEM weight per unit floor area of the segment unit indicates that the unit weights are still relatively higher that the conventional B777 type cylindrical or A380 type elliptic fuselage design. Due to the manufacturing concern associated with multi-bubble fuselage, a Y braced box-type fuselage alternative with special resin-film injected (RFI) stitched carbon composite with foam-core was designed by Boeing under a NASA research contract for the 480 passenger version. It is shown that this configuration can be improved to a modified multi-bubble fuselage which has better stress distribution, for same material and dimension.

Abstract: The invention relates to an aircraft floor (4), preferably a cockpit floor, this floor comprising a plurality of spars (14) running along a longitudinal direction (X) of the aircraft and a plurality of cross-beams (16) assembled to the spars and running along a transverse direction (Y) of the aircraft, the floor also comprising attachment means (64) used to assemble it to fuselage frames. According to the invention, the attachment means comprise a plurality of articulations (66) each connected to one end of one of the cross-beams (16), and enabling rotation about the (X) direction.

Abstract: The unsteady evolution of trailing vortex sheets in ground effect is simulated by the use of a discrete vortex method. The ground effect is included by image method. Two cases of unsteady vortex evolution behind lifting lines (an elliptic loading and a fuselage/flap-wing configuration) are simulated for several ground heights. The present method is validated by comparison of the simulated wake roll-up shapes to published numerical results. Fo ra lifting line with an elliptic loading, the ground has the effect of moving the wingtip vortices laterally outward and suppressing the development of the vortex. An increase in the wing loading has the effect of moving the wingtip vortex more laterally outward. The rotation of the wake vortices behind a fuselage/part-span flap configuration in ground effect is less than the case of flight out of the ground effect. Nomenclature b = span C(�, t) = point on a lifting line in the complex plane C � (� o , t) = point on an image lifting line in the complex plane h = ground height; distance from a lifting line to the ground N = number of point vortices Re =v ortex Reynolds number rc =v ortex core radius

Abstract: Insulation for use in an aircraft fuselage is formed wherein at least one layer (40) made from an open-celled foam (39) that provides acoustic and thermal insulation. The open-celled foam (39) is compression fitted into the airplane fuselage so as to provide effective attachment to the fuselage. The open-celled foam (39) requires minimal attachment treatments. Further, compression fit of the open cell layer (40) is used as an interface around brackets and unrelated hardware to provide superior close-out of gaps that would normally occur using traditional insulation to bracket interfaces. The preferred foam for the open celled compression layer is also relatively moisture resistant (i.e. hydrophobic) in nature and is compressible to between about 0.5 and 10 percent compression, with about 2% compression being ideal for most applications.

Abstract: A rotationally symmetrical anti-collision light (100, 200) for an aircraft utilizing light-emitting diodes (LEDs) is mounted to the fuselage of an aircraft. The LEDs (10, 10', 20, 20') may be configured in one or more concentric rings. The anti-collision light includes a reflector (30, 30') configured to redirect the light emitted by at least one of the rings, so that the light pattern satisfies predetermined specifications.

Abstract: An aircraft propulsion system (10) or a fixed wing aircraft (12) having a power plant or gas turbine engine (18) driving a number of outboard propulsion units (20) mounting to pylons (24) extending laterally outwardly from the fuselage (14) of the aircraft and propelling the fixed wing aircraft (12) via mechanical transmission (28). The propulsion units (20) are powered by the gas turbine engine (18) independently from one another to permit selective shutdown of at least one of the propulsion units (20) while still allowing the remaining propulsion units (20) to be driven by the engine (18). The mechanical transmission (28) is a multibranch transmission wherein an overload decoupling device (48) is provided in each transmission shaft branch (30) set to automatically disconnect an associated propulsion unit (20) when a predetermined critical value is reached. Both the propulsion units (20) and the gas turbine engine (18) having lubrication systems (50, 54) to lubricate transmission components and the decoupling devices (48).

Abstract: Structural analysis and design of efficient pressurized fuselage configurations for the advanced Blended-Wing-Body (BWB) flight vehicle is a challenging problem. Unlike a conventional cylindrical pressurized fuselage, stress level in a box type BWB fuselage is an order of magnitude higher, because internal pressure primarily results in bending stress instead of skin-membrane stress. In addition, resulting deformation of aerodynamic surface could significantly affect performance advantages provided by lifting body. The pressurized composite conformal multi-lobe tanks of X-33 type space vehicle also suffered from similar problem. In the earlier BWB design studies, Vaulted Ribbed Shell (VLRS), Flat Ribbed Shell (FRS); Vaulted shell Honeycomb Core (VLHC) and Flat sandwich shell Honeycomb Core (FLHC) concepts were studied. The flat and vaulted ribbed shell concepts were found most efficient. In a recent study, a set of composite sandwich panel and cross-ribbed panel were analyzed. Optimal values of rib and skin thickness, rib spacing, and panel depth were obtained for minimal weight under stress and buckling constraints. In addition, a set of efficient multi-bubble fuselage (MBF) configuration concept was developed. The special geometric configuration of this concept allows for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls, while the outer-ribbed shell prevents buckling due to external resultant compressive loads. The initial results from these approximate finite element analyses indicate progressively lower maximum stresses and deflections compared to the earlier study. However, a relative comparison of the FEM weight per unit floor area of the segment unit indicates that the unit weights are still relatively higher that the conventional B777 type cylindrical or A380 type elliptic fuselage design. Due to the manufacturing concern associated with multi-bubble fuselage, a Y braced box-type fuselage alternative with special resin-film injected (RFI) stitched carbon composite with foam-core was designed by Boeing under a NASA research contract for the 480 passenger version. It is shown that this configuration can be improved to a modified multi-bubble fuselage which has better stress distribution, for same material and dimension.

Abstract: The HondaJet is an advanced, lightweight, business jet featuring an extra large cabin, high fuel efficiency, and high cruise speed compared to existing small business jets. To achieve the high-performance goals, an over-the-wing engine-mount configuration, a natural-laminar-flow wing, and a natural-laminar-flow fuselage nose were developed through extensive analyses and wind-tunnel tests. The wing is metal, having an integral, machined skin to achieve the smooth upper surface required for natural laminar flow. The fuselage is constructed entirely of composites; the stiffened panels and the Sandwich panels are cocured integrally in an autoclave to reduce weight and cost. The prototype aircraft has been designed and fabricated. Major ground tests such as structural proof tests, control-system proof test, system function tests, and ground-vibration tests have been completed. The first flight was conducted on 3 December 2003, and flight testing is currently underway. The aerodynamic, aeroelastic, structural, and systems designs and tests conducted during the development are described.

Abstract: This paper demonstrates coupled time-domain computational-fluid-dynamics (CFD) and computationalstructural-dynamics simulations for flutter analysis of a real aircraft in the transonic regime. It is shown that a major consideration for a certain class of structural models is the transformation method, which is used to pass information between the fluid and structural grids. The aircraft used for the calculations is the BAE Systems Hawk. A structural model, which has been developed by BAE Systems for simplified linear flutter calculations, only has a requirement for O(10) degrees of freedom. There is a significant mismatch between this and the surface grid on which loads and deflections are defined in the CFD calculation. This paper extends the constant volume tetrahedron tranformation, previously demonstrated for wing-only aeroelastic calculations, to multicomponent, or full aircraft, cases and demonstrates this for the Hawk. A comparison is made with the predictions of a linear flutter code.

Abstract: The frame (1) includes at least one outer flange (2) attached to the window opening edge of the aircraft fuselage, and at least one inner flange (3) and a vertical flange (4) which support a window piece e.g. glass panel. The frame is made from resin reinforced with unidirectionally-arranged fiber bundles. An independent claim is also included for a window frame manufacturing method.

Abstract: A structural frame for a fuselage including a lower arched structure (1), a cross-piece (2), supporting a floor element (3) and a connecting lattice (4) between the lower arch (1) and the cross-piece (2), the connecting lattice (4) more particularly may comprise bars (4a, 4b) the ends of which are spliced to the cross-piece (2) and the lower arch (1).

Abstract: Methodologies used in industry for designing transport aircraft composite fuselage structures are discussed. Several aspects of the design methodologies are based on assumptions from metallic fuselage technology, which requires that full-scale structures be tested with the actual loading conditions to validate the designs. Composite panels that represent crown and side regions of a fuselage structure are designed by the use of this approach and tested in biaxial tension. Descriptions of the state-of-the-art test facilities used for this structural evaluation are presented. These facilities include a pressure-box test machine and a D-box test fixture in a combined loads test machine, which are part of a combined loads test system. Nonlinear analysis results for a reference shell and a stiffened composite panel tested in the pressure-box test machine with and without damage are presented. The analytical and test results are compared to assess the ability of the pressure-box test machine to simulate a shell stress state with and without damage. A combined loads test machine for testing aircraft primary structures is described. This test machine includes a D-box test fixture to accommodate curved stiffened panels, and the design features of this test fixture are presented. Finite element analysis results for a curved panel to be tested in the D-box test fixture are also discussed.

Abstract: In March 2002, a 25-ft/s vertical drop test of a composite fuselage section was conducted onto water. The purpose of the test was to obtain experimental data characterizing the structural response of the fuselage section during water impact for comparison with two previous drop tests that were performed onto a rigid surface and soft soil. For the drop test, the fuselage section was configured with ten 100-lb. lead masses, five per side, that were attached to seat rails mounted to the floor. The fuselage section was raised to a height of 10-ft. and dropped vertically into a 15-ft. diameter pool filled to a depth of 3.5-ft. with water. Approximately 70 channels of data were collected during the drop test at a 10-kHz sampling rate. The test data were used to validate crash simulations of the water impact that were developed using the nonlinear, explicit transient dynamic codes, MSC.Dytran and LS-DYNA. The fuselage structure was modeled using shell and solid elements with a Lagrangian mesh, and the water was modeled with both Eulerian and Lagrangian techniques. The fluid-structure interactions were executed using the fast general coupling in MSC.Dytran and the Arbitrary Lagrange-Euler (ALE) coupling in LS-DYNA. Additionally, the smooth particle hydrodynamics (SPH) meshless Lagrangian technique was used in LS-DYNA to represent the fluid. The simulation results were correlated with the test data to validate the modeling approach. Additional simulation studies were performed to determine how changes in mesh density, mesh uniformity, fluid viscosity, and failure strain influence the test-analysis correlation.

Abstract: The invention relates to a convertible aircraft operating method According to the invention, the aircraft comprises: a fuselage, standard fixed wings which are equipped with ailerons, a tail unit with flight-control surfaces, engines, a rotor with blades, a transmission which is placed between the engines and the rotor and which is equipped with rotor clutch and braking means, a landing gear, means for transition from helicopter mode to gyroplane mode and vice versa, and means for direct or reverse transition from gyroplane/helicopter mode to aeroplane mode The lift for a range of low speeds is produced by means of the rotor, while the lift for a range of high speeds is produced by means of the wings In addition, the lift for a range of intermediate speeds can be produced using the wings and the rotor in gyroplane mode simultaneously, and take-off and landing can be performed in gyroplane mode or in helicopter mode with the engines coupled to the rotor The aircraft comprises a hybrid helicopter/gyroplane/aeroplane aircraft and, as such, can perform the direct or reverse transition to aeroplane mode from both helicopter mode and gyroplane mode

Abstract: The design and attachment of an insulation package according to an exemplary embodiment, which may be installed near the fuselage structure, is believed protect the cabin region of an aircraft against fire whose flames act on the insulation package from outside the aircraft environment, thus clearly facilitating evacuation of the passangers from the vehicle. The insulation package arrangement may comprise several fuselage insulation packages (19-22) of an elongated form. These packages may adjoin the aircraft fuselage structure in the direction of the longitudinal axis (9) of the aircraft. They may longitudinally adjoin a support surface (31a) of the stringers (31) which are attached to the aircraft fuselage, or longitudinally adjoin an inner area (33a) of a panel of outer skin and are attached to both longitudinal sides of the ribs (32). Furthermore those insulation packages (19-22) may be completely enclosed by a burn-through-proof foil (11) which is arranged in a space enclosed by interior paneling and by the panels of the outer skin. The design of a fuselage insulation package may be implemented with burn-through~proof insulation of a larger cross section and/or a burn-through-proof barrier layer of a smaller cross section which are arranged within the fuselage insulation package either singly or in combination. In this arrangement, the insulation or the barrier layer extends near to or adjacent to an interior wall region of the foil wall. As an alternative only that insulation which on the longitudinal end of the fuselage insulation package continues outward with a flat insulation end section is atteched outside of and adjacent to the foil (11) circumference of the fuselage insulation package (19-22). Said end section of insulation is attached to a rib-attachment region which is arranged below the respective longitudinal sides of a rib (32) and near the stringer (31) by means of burn-through-proof attachment elements (4, 13).

Abstract: A numerical simulation of the aircraft impact into the exterior columns of the World Trade Center (WTC) was done using LS-DYNA. For simplification, the fuselage was modeled as a thin-walled cylinder, the wings were modeled as box beams with a fuel pocket, and the engines were represented as rigid cylinders. The exterior columns of the WTC were represented as box beams. Actual masses, material properties and dimensions of the Boeing 767 aircraft and the exterior columns of the WTC were used in this analysis. It was found that about 46% of the initial kinetic energy of the aircraft was used to damage columns. The minimum impact velocity of the aircraft to just penetrate the exterior columns would be 130 m∕s . It was also found that a Boeing 767 traveling at top speed would not penetrate exterior columns of the WTC if the columns were thicker than 20 mm .

Abstract: The literature on turbulent boundary layer pressure fluctuations provides several empirical models which were compared to the measured TU-144 data. The Efimtsov model showed the best agreement. Adjustments were made to improve its agreement further, consisting of the addition of a broad band peak in the mid frequencies, and a minor modification to the high frequency rolloff. The adjusted Efimtsov predicted and measured results are compared for both subsonic and supersonic flight conditions. Measurements in the forward and middle portions of the fuselage have better agreement with the model than those from the aft portion. For High Speed Civil Transport supersonic cruise, interior levels predicted by use of this model are expected to increase by 1-3 dB due to the adjustments to the Efimtsov model. The space-time cross-correlations and cross-spectra of the fluctuating surface pressure were also investigated. This analysis is an important ingredient in structural acoustic models of aircraft interior noise. Once again the measured data were compared to the predicted levels from the Efimtsov model.

Abstract: The surface and interior response of a Cessna Citation fuselage section under three different forcing functions (10–1000 Hz) is evaluated through spatially dense scanning measurements. Spatial Fourier analysis reveals that a point force applied to the stiffener grid provides a rich wavenumber response over a broad frequency range. The surface motion data show global structural modes (≲150 Hz), superposition of global and local intrapanel responses (∼150–450 Hz), and intrapanel motion alone (≳450 Hz). Some evidence of Bloch wave motion is observed, revealing classical stop/pass bands associated with stiffener periodicity. The interior response (≲150 Hz) is dominated by global structural modes that force the interior cavity. Local intrapanel responses (≳150 Hz) of the fuselage provide a broadband volume velocity source that strongly excites a high density of interior modes. Mode coupling between the structural response and the interior modes appears to be negligible due to a lack of frequency proximity and ...

Abstract: The temperature distribution of the exposed rear fuselage of fighter aircraft, which is necessary for estimating passively emitted infrared signature levels, is predicted. Hot combustion gases flowing through the gas-turbine engine and aerodynamic heating of the rear fuselage at high freestream Mach numbers are the two heat sources. Turbulent forced convection and radiation are the dominant heat-transfer modes, and hence two thermal models are developed: only turbulent forced convection, and combined turbulent forced convection and radiation. Variation in transport and thermophysical properties of fluids is also modeled. Gas-dynamic equations are used to obtain local flow parameters, considering the role of friction, heat transfer, and area variation. A radiation-shape-factor analysis for the complete layout is presented that accounts for mutual, partial, or complete visibility of emitter and receiver elements. A closed-form solution for shape factors between ring elements is derived for the annulus. The error in temperature estimation by neglecting parameters such as aerodynamic heating, variation in transport, and thermophysical and local-flow properties of fluid is found to be significant. We conclude with a discussion of the role of altitude and flight Mach number.

Abstract: An air-launched aircraft (10) includes deployable wings (16, 18), elevons (20, 22), and vertical fins (26, 28) that deploy from a fuselage (12) during flight. The aircraft may include a control system for operating the elevons, a communication system, and batteries for powering the systems. In addition, the aircraft may include a payload module (14) that mates with an interface in the fuselage. The payload module may include any of a variety of payloads, including cameras, sensors, and/or radar emitters. The aircraft may be powered or unpowered, and may be very small, for example, less than on the order of 10 kg (22 pounds). The deployable surfaces of the aircraft may be configured to deploy in a pre-determined order, allowing the aircraft automatically to enter controlled flight after being launched in a tumbling mode.