Abstract: An adaptive aircraft tail assembly includes a tail boom, and a tail structure. The tail structure has a plurality of control surfaces configured to alter or maintain flight characteristics of the aircraft. The tail structure also has at least (i) one or more photovoltaic cells, or (ii) at least one or more solar thermal collection cells, or (iii) both photovoltaic cells and solar thermal collection cells. The tail structure also includes a rotational pivot configured to rotationally attach the tail structure to the tail boom. Preferably, the plurality of control surfaces are configured so as to be manipulatable to rotate the tail structure about a central axis of the tail boom via the rotational pivot.

Abstract: A stabilizing and directional-control surface of an aircraft includes a vertical stabilizer and a rudder that deflects relative to the vertical stabilizer. The rudder includes an internal profile that is extendable and retractable by an actuating system. An aerodynamic control surface area of the rudder is increased when the internal profile of the rudder is extended as compared to the aerodynamic control surface area of the rudder when the internal profile of the rudder is retracted.

Abstract: An apparatus comprises a structure having a first side, a second side substantially opposite to the first side, a flexible skin, and a plurality of deformable assemblies. The flexible skin is attached to the first side and the second side of the structure. The plurality of deformable assemblies is moveably connected to the structure, in which each deformable assembly in the plurality of deformable assemblies has a vertex and a base. The each deformable assembly in the plurality of deformable assemblies has a height that is capable of changing to change a shape of the structure and a shape of the flexible skin.

Abstract: The invention comprises varying the ratio between the control surface local chord (I) and the vertical stabilizer local chord (L) along the height of the vertical tail (2) in order to adapt the local value of the coefficient of the side lift applied to the vertical tail (2) to a maximum acceptable value of the side lift coefficient.

Abstract: An aerodynamic body of an aircraft with an air outlet opening and an air intake opening that communicates with the air outlet opening via an air conduit is described. A flow delivery driver device for influencing the flow within the air conduit is integrated into the air conduit. The surfaces of the aerodynamic body in the body chord direction include at least one air outlet opening in the front region of the aerodynamic body, and at least one air intake opening on the upper surface of the aerodynamic body and in the rear region of the aerodynamic body and/or on the upper surface of the aerodynamic body in the trailing edge region and/or on the lower surface of the aerodynamic body in the trailing edge region. Arrangements of a main wing and an adjustable flap, and an aircraft with such an aerodynamic body are also described.

Abstract: A tail section for an aerospace vehicle is provided. The tail section comprises a rudder which is movable about an axis to generate a yawing moment on the aerospace vehicle. The tail section further comprises a thruster having, in flow series, an air intake, an electrically powered device for accelerating the air received through the intake, and an air outlet which directs the accelerated air to increase the yawing moment generated by the rudder.

Abstract: A lateral force joint, with which a plate-like spar can be fastened to a fuselage strap mounted on an airplane fuselage. The plate-like spar has a through bore, into which a fastening cylinder mounted on the flange can be inserted for connecting both components. In order to allow for the location of the spar to be adjusted with respect to the fuselage strap, a bearing bush of the spar is eccentrically fitted into the through bore. Correspondingly, the fastening cylinder of the fuselage strap is mounted eccentrically on a bearing bush so that a rotation of the bearing bushes allows the lateral force joint to be adjusted.

Abstract: A lifting or stabilising component (11) for an aircraft comprising in the area of its trailing edge a rotary control element (13) rotating around an axis (61) with at least one slot (21) between the tip (15) of the component (11) and the control element (13), the edges of which are configured such that the distance therebetween, that is, the dimension of the slot (21), is constant for different angles of deflection of the control element (13), being located between them a seal (23) that assure the aerodynamic continuity of the component (11) when the control element (13) is at rest. In a preferred embodiment, the component (11) is a horizontal tail stabilizer and the control element (13) is a rudder.

Abstract: An improvement and method for improved structural dynamic stability for 20 and 30 Series Learjets are disclosed. The improvement includes a redistribution of the elevator mass balance to uncouple the elevator rotational motion from the stabilizer translation motion for the higher order horizontal frequencies having node lines in the proximity of the mass outboard counterbalance weights. The original tail section includes a rudder, and a horizontal stabilizer supporting an elevator mounted adjacent the rudder. The elevator includes a proximal end adjacent the rudder and a distal end that includes a counterbalance portion. The improvement includes replacement of an original mass counterbalance weight from within the counterbalance portion with a new mass counterbalance weight of less mass, and the inclusion of additional mass counterbalance weights disposed within the elevator and interposed between the proximal end and the counterbalance portion.

Abstract: A redundant electromechanical actuator is provided that includes first and second electric motors, a common output shaft, first and second drive cable drums, first and second driven cable drums, a first cable, and a second cable. The first and second electric motors are each adapted to be controllably energized to supply a drive torque. The common output shaft is coupled to receive the drive torque supplied from one or both of the motors. The first and second drive cable drums are coupled to the common output shaft to receive the drive torque from one or both of the motors. The first and second driven cable drums are each configured, upon being driven, to rotate. The first cable is coupled between the first drive cable drum and the first driven cable drum, and the second cable is coupled between the second drive cable drum and the second driven cable drum.

Abstract: A rudder of a commercial aircraft is provided that is divided along its longitudinal direction in at least one region and the parts of the rudder can be spread against the air flow surrounding the aircraft by means of an actuator in order to decelerate the aircraft.

Abstract: The invention relates to a control surface (3) for an aerodynamic lifting surface (2) of aircraft comprising a main closing rib (9) located at one end of the control surface (3) to which a main torsion bar (8) is joined, the mentioned torsion bar (8) being joined at its other end to a lever system (14) on which at least one actuator (15) acts, such that it is possible to act on the rotation of the mentioned control surface (3) during the flight of the aircraft. The invention also relates to an actuation method for actuating such a control surface (3).

Abstract: A large-capacity airplane principally includes a fuselage which has no region of constant width between the front and the rear and a wing fixed to the fuselage. As a preference, the engines are fixed at the rear under a horizontal tail held above the fuselage by vertical stabilizers and maintenance wells are formed in the fuselage in vertical alignment with each engine to allow the engines to be fitted and removed using conventional means. The width of the fuselage is determined, on the one hand, so that the airplane landing gear is fixed to the fuselage and in the up position is included within the interior volume of the fuselage and, on the other hand, so that the rear engines are above the fuselage in order to make use of the beneficial effects of this position.

Abstract: A pivoting coupling system of a large dihedral empennage to the tail fuselage of an aircraft, in which the empennage includes a right lateral box and a left lateral box arranged at a large dihedral angle in the tail fuselage of the aircraft. The pivoting coupling system including a horizontal central box joining the lateral boxes, and which includes a rear spar; and a linkage for horizontally linking the central box to a frame structure of the tail fuselage of the aircraft, which allow the empennage to rotate vertically about a horizontal linkage shaft between a negative maximum angle of incidence and a positive maximum angle of incidence in response to the actuation of an actuator which is connected to the empennage and to a structural element of the fuselage of the aircraft.

Abstract: Rudder assist mechanisms and methods are capable of being operably connected to an aircraft's rudder control system. The rudder assist mechanisms most preferably have over-the-center spring biasing functions so as to cause either substantially no spring force (i.e., when the linkage is over the spring-bias center) or substantially all spring force (i.e., when the linkage is left or right of the spring-bias center) to be exerted on the rudder control system. The rudder assist mechanism may include a control spring assembly, and a linkage assembly which operably connects the control spring assembly to the rudder control assembly. The linkage assembly is moveable operably between a null position wherein substantially no spring force of the control spring assembly is transferred to the rudder control system by the linkage assembly, and right and left spring-biased positions wherein right and left spring forces of the control spring assembly are transferred to the rudder control system, respectively.

Abstract: The invention relates to a flight control system for an aircraft comprising control surfaces (111, 211, 331, 431, 541, 121-225), and actuators associated with said control surfaces to control flight functions of roll, yaw, pitching and aerodynamic braking of the aircraft. All of the actuators (111a1-211a2, 121a-225a) associated with the control surfaces controlling at least one of said flight functions are electromechanical actuators. A part of said control surfaces associated with the electromechanical actuators are divided control surfaces (111, 211, 331, 431, 541), each of said divided control surfaces being composed of at least two independent surfaces (113, 114, . . .

Abstract: An aircraft (1) horizontal stabiliser (2) which can vary its sweep angle (6) by rotating around an essentially vertical axis (4) with respect to the direction of flight of the aircraft (1) when the aircraft (1) is in flight at high speeds, close to the speed of sound, this axis (4) being contained in the plane of symmetry of the aforementioned aircraft (1), the aforementioned horizontal stabiliser (2) being is a single and structurally continuous component such that it does not transmit a bending moment to the fuselage (14) structure of the aircraft (1), which permits it to be of reduced weight, such that the horizontal stabiliser (2) rotates in a single way and a single direction around the axis (4) to achieve the sweep angle (6) required for high-speed flight.

Abstract: An aircraft has at least two rudder units in a non-central arrangement. The rudder units have a profile and are rigidly secured to the aircraft. The profile is adjusted in such a way as to yield an inflow at a positive geometric angle of incidence, so that the profile generates a lift force having a component pointing in the flight direction. This makes it possible to offset a portion of the aircraft drag.

Abstract: An attachment system for attaching a tail unit to an attachment surface of an aircraft is disclosed. An attachment surface may be a fuselage, for example. In an arrangement, the attachment system includes a mounting with a first bearing surface that is capable of resting against the tail unit, and a second bearing surface that is capable of resting against the attachment surface. The first bearing surface and the second bearing surface, of which there is at least one, include a common line of contact. In the arrangement, the planes of the first bearing surface and of the second bearing surface are at an angle that differs from 0° and 180°.

Abstract: A vertical tail for use with an aircraft or other form of mobile platform. The vertical tail includes a main element which is fixedly secured to the mobile platform, and a leading edge element that is movably secured to the main element. The cross section of the leading edge element is symmetric about the cruise chord line of the tail. The leading edge element can be pivoted and/or extended to create a gap with the main (fixed) element. The movable leading edge element is used to increase the maximum yawing moment provided by the vertical tail. The maximum yawing moment is increased when air flow is incident from either side of the vertical tail.

Abstract: An active winglet includes a body portion substantially parallel to a wing of an aircraft, as if it were an extension of the wing, and an angled portion that projects at an upward angle from the body portion. The body portion is attachable to an aircraft wing and includes a controllable airflow modification device coupled thereto. By virtue of having a controllable airflow modification device, the winglet is capable of adjusting a control surface of the controllable airflow modification device in response to in-flight conditions, to reduce wing loads, increase range, and increase efficiency.

Abstract: An aircraft tail assembly planform comprising curvilinear leading edges (21) and trailing edges (22), with an aircraft tail assembly configuration in which the hinge line (13) is rectilinear and has a non-constant percentage with respect to the chord (50) in each section (51), being the front (11) and rear (12) spars rectilinear with a non-constant percentage with respect to the chord (50) in each section (51), or being these spars (11, 12) curvilinear with a constant or non-constant percentage with respect to the chord (50) in each section (51).

Abstract: A roll steering method subdivides the direction of an aircraft control surface into two elements and, during a roll control operation using ailerons, steers an upper element of the control surface in the roll direction and a lower element in the opposite direction.

Abstract: A system and method for assembling and operating a solar powered aircraft, composed of one or more modular constituent wing panels. Each wing panel includes at least one hinge interface that is configured to rotationally interface with a complementary hinge interface on another wing panel. When a first and second wing panel are coupled together via the rotational interface, they can rotate with respect to each other within a predetermined angular range. The aircraft further comprises a control system that is configured to acquire aircraft operating information and atmospheric information and use the same alter the angle between the wing panels, even if there are multiple wing panels. One or more of the wing panels can include photovoltaic cells and/or solar thermal cells to convert solar radiation energy or solar heat energy into electricity, that can be used to power electric motors.

Abstract: The angular profile consists of sections (4) arranged one after another joined to each other, in order to simplify the connections between said sections (4). It is characterized in that the adjacent ends of each section (4) present in one of their sides an extension (7) complementary to a recess (8) provided in the same side of the end of the adjacent section (4) for locating the extension (7) in the recess (8) and, by means of bolted joints (9), to fix the extension (7) in the opposite side of the adjacent section (4) constituting the profile. The sections (4) and the bolted joints (9) are made of metal to provide electrical continuity. The angular profile constitutes the trailing edge of a tail rudder (1) of an HTP of an aircraft.

Abstract: The invention relates to a fairing system for a horizontal stabiliser of an aircraft and to a process for installing said system, such that the fairing system for the sealing of the horizontal stabiliser (2) and the fuselage (1) of the aircraft comprises a primary fairing (50, 51) arranged on the horizontal stabiliser (2) of the aircraft, the mentioned primary fairing (50, 51) being solidly connected to the mentioned horizontal stabiliser (2) which is in turn trimmable with respect to the fuselage (1), the fairing system in turn comprising a secondary fairing (80, 81) such that the primary fairing (50, 51) is coupled to the horizontal stabiliser (2) through supports (30) arranged in the direction of the chord of the primary fairing (50, 51), said supports (30) being located in the outer part of said primary fairing (50, 51) with respect to the fuselage (1), the secondary fairing (80, 81) being arranged in the outer part of the supports (30), the purpose of said secondary fairing (80, 81) being to cover the me

Abstract: An aircraft structure is provided. The aircraft structure for an aircraft includes a fuselage, front wings, rear small wings, vertical winglets, and four levels. The fuselage has a cross-section of substantially a half circle fuselage shape. Being is wide enough to provide lifting force, and includes four levels separated by multi partition structure. The front wings are disposed horizontally in front portions of the fuselage. The rear small wings are disposed horizontally in rear portions of the fuselage. The vertical winglets are disposed at the wingtips of the rear small wings. The first level disposed at a bottom of the four levels includes a fuel tank storages and a plurality of landing gear bays. The second level disposed at a middle of the four levels includes cargo bay. The third level may comprise a top cockpit and a plurality of passenger cabins. The fuselage provides major portion of lifting force and the wings provides steering force, and lifting force.

Abstract: Trailing edge (3) of an aircraft stabilizer surface (1), where this surface (1) is manufactured of a composite material and comprises an outer cladding (40) and an inner cladding (41) that are connected by a connecting clip type element (20) on this trailing edge (3), the connecting clip type element (20) comprising at its ends some recesses (23) used for coupling to the inner zone of the upper and lower claddings (40, 41) of the stabilizer surface (1), such that the connecting clip type element (20) is flexible enough to be pinched so that its ends will be housed, by means of these recesses (23), between the outer and inner claddings (40, 41) of the aircraft stabilizer surface (1), whereby the outer zone of the stabilizer surface (1), on its trailing edge (3), is formed by a continuous aerodynamic surface without changes of gradient.

Abstract: A monomolecular carbon-based film can be placed on an aircraft part, such as the leading edge designed to directly impinge against air during flight, ascent or descent, in order to form a smooth surface having increased lubricity and reduced air friction. The aircraft part may be in the form of a helicopter rotor, wing, propeller, fin, aileron, nose cone, and the like. The monomolecular carbon-based film can be deposited on the aircraft part, for example, using a reactor that includes a bed of silica and through which emissions from a diesel engine are passed. The monomolecular carbon-based film decreases air friction and increased lift of a modified aircraft that includes an aircraft part treated with the film.

Abstract: A lifting or stabilising component (11) for an aircraft comprising in the area of its trailing edge a rotary control element (13) rotating around an axis (61) with at least one slot (21) between the tip (15) of the component (11) and the control element (13), the edges of which are configured such that the distance therebetween, that is, the dimension of the slot (21), is constant for different angles of deflection of the control element (13), being located between them sealing means (23) that assure the aerodynamic continuity of the component (11) when said control element (13) is at rest. In a preferred embodiment, the component (11) is a horizontal tail stabilizer and the control element (13) is a rudder.

Abstract: Aircraft horizontal stabilizer surface (8) in which the sweep angle (40) of this surface (8), where this angle (40) is the one formed by the projection of the reference line of points located at 25% of the local chord (19) of the horizontal stabilizer surface (8) on a plane perpendicular to the aircraft plane of symmetry (21), and which also contains this plane to the flight direction of the aircraft with respect to the aircraft plane of symmetry (21), is less than 90 degrees, with this angle (40) being measured in the flight direction of the aircraft. In addition, the structural connection of this horizontal stabilizer surface (8) to the aircraft fuselage (1) is located at a closing frame (13) of this fuselage (1).

Abstract: A rudder of a commercial aircraft is provided that is divided along its longitudinal direction in at least one region and the parts of the rudder can be spread against the air flow surrounding the aircraft by means of an actuator in order to decelerate the aircraft.

Abstract: An aircraft includes a fuselage having a shape elongated along a longitudinal axis X of the aircraft and at least one wing fixed to the fuselage between the front end and the rear end of the fuselage. The fuselage includes a substantially cylindrical central part and a rear tapered part on which a vertical empennage is fixed. Between a section connecting the rear part with the central part of the fuselage and the rear end the maximum width of each section of the fuselage is constant or increasing rearwards up to a maximum width L of the fuselage, the height of each section of the fuselage is decreasing rearwards in the direction of the negative X, so that the rear end of the fuselage forms a trailing edge having a small thickness which is substantially horizontal in an aircraft reference system and substantially rectilinear.

Abstract: The present invention relates to a sandwich-structure flat active actuator that is symmetrical relative to the mid-plane comprising at least one piezoelectric flat layer having an active direction superpose and fixedly attached to at least one passive flat layer of fabric with rigid warp and weft oriented in two directions forming a mesh; the two directions of each layer of fabric are the same. The active direction of each piezoelectric layer is oriented along one and the same diagonal of the mesh of the layers of fabric. The invention is also aimed at a device consisting of two actuators mounted head-to-tail. The invention applies to the active torsion of structures such as, for example, a helicopter blade.

Abstract: A rudder pedal mechanism is configured with foreign object strike tolerance and/or with an articulating brake. The foreign object strike tolerance is implemented via a force transfer mechanism that is responsive to an input force supplied to the rudder pedal. The force transfer mechanism is configured to supply a transfer force to a rudder position command unit when the input force is supplied in a first direction, and to not supply the transfer force to the rudder position command unit when the input force is supplied in the second direction and exceeds at least a predetermined force magnitude. The articulating brake is implemented via a brake rod assembly that is rotationally coupled to the rudder pedal, and includes a first rod that is rotationally coupled to the rudder pedal, and a second rod that surrounds at least a portion of the first rod.

Abstract: The invention provides an aircraft, comprising a fuselage (8), an airfoil (1) mounted to the fuselage (8) and a flap (2; 3) for steering the aircraft. Furthermore, connecting means (18; 79) articularly connect the flap (2; 3) to the airfoil (1) such that the flap (2; 3) is allowed to rotate around a rotation axis (28; 82) substantial parallel to the trailing or leading edge (4; 5) of the airfoil (1) between a retracted position (I) and an extended position (II) and to translate in a direction (54a; 98) substantially parallel to the rotation axis (28; 82). A rod (43; 95) articularly connect the flap (2; 3) to the airfoil (1) or to the fuselage (8), wherein the rod (43; 95) defines the translation of the flap (2; 3) in the direction (55a; 98) parallel to the rotation axis (28; 82). Hence, by way of the invention, forces acting of the flap (2; 3) in a direction (55a; 98) parallel to the rotation axis (28; 82) can be taken up by the rod (43; 95).

Abstract: A bi-directional flight control surface mechanism comprises a fixed part V of a vertical stabiliser having leading and trailing edges and a rudder having leading and trailing edges. The rudder is mounted relative to the fixed part of the vertical stabiliser so that the trailing edge of the fixed structure is adjacent the leading edge of the rudder. The rudder is movable from a neutral position in which the rudder lies in line with the fixed part of the vertical stabiliser to define a notional centreline to first and second angled positions in which the rudder is angled relative to the fixed part of the vertical stabiliser at a non-zero angle on one or other side of the notional centreline. An actuator is provided to drive the rudder relative to the fixed part of the vertical stabiliser so as to vary the distance between said part of the trailing edge of the fixed part of the vertical stabiliser and the leading edge of the rudder.

Abstract: A system that acts on a rudder (L) having its stop parts commanded or moved by an actuator (A) is described. The rudder system comprises a reversible mechanical system with movable stop parts working based on control/command signals, having stop parts (M, M?) inside of the field of movement of the actuator (A), interfering in the movement of said actuator; the mentioned control signals are based on informations of the aircraft such as traction of the engines or other signal from each engine (pressure, temperature or axis rotation speed); flight speed (“airspeed”); altitude; skidding angle; aircraft on the ground.

Abstract: A method for controlling control surfaces. A position limit is identified for movement of a control surface based on a load limit set for the control surface and a number of vehicle current operation parameters to form an identified position limit. Responsive to receiving a command to move the control surface on a vehicle to a new position, the control surface is commanded to move to a position within the identified position limit.

Abstract: An airborne vehicle having a wing-body which defines a wing-body axis and appears substantially annular when viewed along the wing-body axis, the interior of the annulus defining a duct which is open at both ends. A propulsion system is provided comprising one or more pairs of propulsion devices, each pair comprising a first propulsion device mounted to the wing-body and positioned on a first side of a plane including the wing-body axis, and a second propulsion device mounted to the wing-body and positioned on a second side of the plane including the wing-body axis. A direction of thrust of the first propulsion device can be adjusted independently of the direction of thrust of the second propulsion device and/or a magnitude of thrust of the first propulsion device can be adjusted independently of the magnitude of thrust of the second propulsion device.

Abstract: According to the invention, the horizontal stabilizer comprises a fixed intermediate part secured to the structure of said aircraft and at least one mobile edge part able to slide with respect to said fixed intermediate part, transversely to the length thereof. Mobile flaps are able to provide the aerodynamic continuity of the suction face and of the pressure face of said stabilizer when said mobile edge part is deployed.

Abstract: An aircraft has a vertical fin fastened to the rear and above a fuselage of elongated form, essentially in a vertical plane of symmetry of the aircraft. The vertical fin has at least two stable positions, an extended position and a returned position, such that a surface of the vertical fin, subjected to an aerodynamic flow when the aircraft is in flight, is modified in position or in surface between the returned position and the extended position, so that the aerodynamic drag of the vertical fin is reduced in the returned position under given flight conditions compared to the extended position. The change from one surface to another of the vertical fin is accomplished by modifying the geometry of the vertical fin or by displacing the vertical fin relative to the fuselage so that the vertical fin, for example, is more or less inside the fuselage, or more or less immersed in the wake zone of the fuselage in which the local dynamic pressure Pd is reduced relative to the infinitely upstream dynamic pressure Pd0.

Abstract: A trimmable horizontal stabilizer is provided adjacent to the fuselage of an aircraft and has a predetermined aerodynamic profile. A movable elevator is arranged adjacent to the horizontal stabilizer. The horizontal stabilizer has a load-bearing structure which extends in the span direction and is firmly connected to the fuselage of the aircraft, and movable areas which are connected to the load-bearing structure such that they can move and can be moved independently of the elevator in order to trim the horizontal stabilizer by varying the aerodynamic profile.

Abstract: An aircraft attitude control configuration enables control surfaces to provide attitude control for an aircraft at hover or low air speed conditions. The aircraft attitude control configuration includes a plurality of thrusters mounted to an aircraft for thrusting air, a first control surface kinematically coupled to the aircraft at a position downstream of a first thruster to enable a first vector force to be generated by a portion of the thrusted air from the first thruster on the first control surface, and a second control surface kinematically coupled to the aircraft at a position downstream of a second thruster, the first and the second control surfaces being displaced symmetrically on opposite sides of a longitudinal axis of the aircraft, the second control surface being configured to be independently and differentially movable with respect to the first control surface to enable a second vector force to be generated by a portion of the thrusted air from the second thruster on the second control surface.

Abstract: Stabilizing and directional-control surface of an aircraft, said surface comprising a vertical stabilizer (2) and a rudder (3), it being possible for said rudder (3) to be deflected relative to the vertical stabilizer (2), and moreover the rudder (3) comprises an internal profile (10) that can be extended and retracted by means of an actuating system (40) relative to the rest of the structure of the rudder (3), so that the stabilizing and control surface, in the retracted position of the internal profile (10) of the rudder (3), is an excellent aerodynamic surface in normal flying conditions, and at the same time increase of the aerodynamic control surface of the vertical stabilizer (2) is achieved for requirements of controllability of the aircraft at low speeds of said aircraft and against strong yawing moments acting thereon, in the position in which the internal profile (10) of the rudder (3) is extended, and moreover the structure of the rudder (3) can be opened, to permit extension of the internal profil

Abstract: The invention concerns a method which consists in subdividing the direction of the control surface into at least two elements (6I, 6S) and, during a roll control using the ailerons (5G, 5D), steering the upper element (6S) in the roll direction and the lower element (6I) in the opposite direction.

Abstract: A fan control effector for an aircraft comprises at least one blade configured to be pivotably deployable in a radially outward direction from a retracted position to a deployed position such that the blade extends out of the aircraft. The fan control effector may be mounted in a symmetrical arrangement about a longitudinal axis on opposing wings of the aircraft. Furthermore, the fan control effector may comprise any number of blades for independent deployment outwardly from the wing. The blades are configured to be angularly deployable along a direction that is non-parallel to the longitudinal axis of the aircraft. The blades may be configured to be deployable sequentially from the wing starting with an initial deployment of an aft-most one of the blades.

Abstract: Substantial reduction in aircraft stowage space can be achieved by stacking aircraft using minimal or no folding of parts of the aircraft. This results in reduced complexity and reduced cost compared to other aircraft compact stowage schemes including for example fully folding an aircraft. Landing gear variable height, length, or orientation can be used to orient all or a portion of the aircraft for stowage. Alternately or in combination with such orientation by landing gear, other portions of the aircraft including tilting rotors and tilting nacelles can be oriented to achieve a geometric configuration conducive to compact stowage. In preferred embodiments, aircraft are moved relative to each other so that a portion of one aircraft is offset relative to a corresponding portion of the other. In especially preferred configurations, a portion of the first aircraft will overlap or lie on top of without touching a portion of the second aircraft.

Abstract: Rotation fitting secured to a frame (6) of the fuselage of an aircraft and linked to a receiving element secured to the empennage (5) which comprises a central fitting (1) and two side fittings (2,3), the central fitting (1) presenting two lugs (1d,1k) with an opening (1e,1j) each and four side wings (1g,1f,1i,1h), and each one of the side fittings (2,3) presenting a lug (2d,3d) with an opening (2e,3e) each and two side wings (2g,2f,3g,3f), the central fitting (1) being secured to a surface of the frame (6) and the side fittings (2,3) in an opposite surface of the frame (6), such that the frame (6) is arranged between the central fitting (1) and the two side fittings (2,3), such that the lugs (1d,1k) of the central fitting (1) make contact with the lugs (2d,3d) of the side fittings, with all the openings (1e,1j,2e,3e) being aligned in order to house a pin (4) associated with the receiving element secured to the empennage (5).

Abstract: An extra narrow wing system for airplanes combining the advantages of sufficient lift like a conventional airplane wing with reduced space requirements. A V-shaped nudge at the back side of the main wing helps to overcome the limitations of extreme delta-wings and allows very narrow wings with angles of as low as 0 degrees but sufficient lift for a propeller or jet operated plane. Front and tail wings are one aerodynamic system. The system results in very narrow airplane designs useful when operating under space constraints.