Abstract: An aircraft wing useful for reducing wing-generated noise comprises: (a) a main element having a leading edge and a trailing edge; (b) at least one slat having a leading edge and a trailing edge, wherein the slat acts cooperatively with the main element to provide lift to the aircraft, and the main edge leading edge and slat trailing edge define a cavity there between; and (c) at least one sound-absorbing structure located on at least a portion of the main element leading edge, at least a portion of the slat trailing edge, in the cavity, or combinations thereof. If located on the main element leading edge or slat-trailing edge, the sound-absorbing structure may be a honeycomb load-bearing structure such as a honeycomb load-bearing structure having a perforated skin or a machine grid structure having an outer skin which is bonded or perforated. If located in the cavity, the sound-absorbing structure may be a flow blocker.

Abstract: A flap support mechanism for attachment of a flap to a rotor blade so as to permit angular displacement with respect thereto has a tension-torsion member disposed coaxial with a hinge axis of the flap. One end at the blade root side of the tension-torsion member is secured to the rotor blade. The other end at the blade tip side of the tension-torsion member is secured to the flap. The tension-torsion member is placed such that it passes through a narrow cavity formed at the interior of the rotor blade. Such construction permits smooth flap motion in spite of a large centrifugal force.

Abstract: A control system for providing attitude control in spacecraft. The control system comprising a primary attitude reference system, a secondary attitude reference system, and a hyper-complex number differencing system. The hyper-complex number differencing system is connectable to the primary attitude reference system and the secondary attitude reference system.

Abstract: An auxiliary power unit for an aircraft includes a gas turbine with a combustion chamber a compressor and a turbine. A compressor is coupled with the turbine, for generating compressed air. A fuel cell (BZ) is present for generating electric energy.

Abstract: An airborne power control system for automatically controlling the power of an aircraft during landing is disclssed. The system includes a computer and a minimum airspeed program as a function of altitude. A radio altimeter or the like senses the instantaneous altitude of the aircraft while a pitot tube or the like measures indicated airspeed. A computer and program compare programmed airspeed with actual airspeed for a given altitude. And, a servomotor is provided for decreasing engine thrust where the actual airspeed exceeds the programmed airspeed at any given altitude. An inhibitor inhibits the decrease in engine thrust if the airspeed drops below the programmed airspeed.

Abstract: Pyrotechnic actuator cover/fin release system designed for high G forces of gun-launched-systems. As safety feature, system allows the pyrotechnic actuator to be installed as one of the last steps of the final testing sequence and removed safely on disassembly. A first cam action develops initial ejection. Subsequently firing the pyrotechnic pushes both the fins and the covers outward via a second cam surface. The release system both holds covers securely latched until commanded to release and then unlatches covers prior to pushing them open. The cover eject sequence results in the flight forward edge of cover being kicked into air stream. Thereafter the cover is lifted and rotates about its aft release hinge until, at about 30 degrees of rotation, the hinge disengages. The fins then are driven into the fully deployed state by a spring and piston/wedge mechanism on the actuator shaft.

Abstract: A helicopter drip pan apparatus for covering and sealing a structural opening in a helicopter. The drip pan apparatus includes a frame member adapted to mount to the structural opening of the helicopter. The frame member has an inwardly-facing peripheral surface extending around the frame member. The frame member and thus the surface have at least one linear portion. A drip pan is selectively affixed to the frame member. The drip pan includes an outwardly-facing peripheral groove which extends around the outside edge of the drip pan. Like the surface of the frame member, the groove has at least one linear portion. The perimeter of the drip pan conforms to the surface of the frame member. A seal member is disposed in the groove of the drip pan to sealingly engaging the surface of the frame member against fluid leakage therethrough.

Abstract: A reusable spacecraft system having two substantially identically reusable return flight space vehicles, one of which is preferably a booster and the other an orbiter, each of which have identical flight control and propulsion systems and have identical but selectably installable components, such as thermal protection disposed on the orbiter but not on the booster. Each vehicle further includes identical payload bays for receiving mission specific payloads, that may be for example, a propellant tankage payload for the booster and a manned mission payload for the orbiter. The use of identical booster and orbiter vehicles reduces costs of development, production and reusable missions.

Abstract: The invention provides a momentum wheel energy storage device for a spacecraft or satellite. The device has six wheels whose axes of rotation are parallel to edges of a tetrahedron. It thus provides redundancy which makes it possible simultaneously to store energy and to generate an arbitrary net angular momentum, even in the event of one of the wheels failing. The failure of one wheel can be accommodated without degrading operation, with energy storage being optimal regardless of the angular momentum to be generated. The failure of two wheels can be accommodated but operation may then be in degraded mode. The invention makes it possible to limit the number of momentum wheels used in a device that can accommodate failures.

Abstract: In an aircraft using fly-by-wire technology, the flight control functions have been integrated into the integrated modular avionics (“IMA”). The new flight control module (“FCM”) resides on the same data bus as the other modules in the IMA and receives power from the same power supply. In addition, the FCM is also connected to a separate power supply to add redundancy to the system. Several benefits arise from this configuration of an FCM. There is no longer a separate chassis needed for the flight control functions, thus resulting in a reduction in weight. In addition, the FCM now has access to all of the data on the IMA bus, instead of a limited amount of data over an ARINC 629 bus. The FCM provides augmentation signals to the actuator control electronics (“ACE”) to aid in the flying of the aircraft. In the event of a failure of the FCM, the ACE still provides enough control to fly the airplane.

Abstract: A solar wing control system for avoiding thermal shock to a solar wing includes an eclipse exit slew profile generator having an eclipse exit slew rate output for rotating the solar wing from an eclipse exit angle to a solar power generation angle to control solar wing heating rate.

Abstract: An electric thrust reverser actuation system includes an electric motors that is controlled to operate in both a motoring mode and a generating mode. During normal actuation system operations, the electric motor is controlled to operate in the motoring mode to move one or more moveable thrust reverser components. During a power interrupt event, in which the motor's primary power source is lost, the motor is controlled to operate in the generating mode. The electrical power generated by the motor during its operation in the generating mode maintains the thrust reverser system locking mechanisms in an energized, unlocked condition, thereby preventing damage to system components.

Abstract: A deicer (10) for a radome (12) includes an active portion (16) through which transmission occurs and a non-active portion (18) which is attached to an aircraft (14). The deicer (10) comprises a cap portion (20) covering the active portion (16) of the radome (12) but unattached thereto. The cap portion (20) comprises layers (36, 38) defining inflatable/deflatable chambers (24), the layers both being made of non-extensible material such as square woven nylon fabric. When the chambers (24) are inflated, the inner layer (38) lifts off of the active portion (16) of the radome (12). The deicer (10) can additionally comprise a skirt portion (22) which is stretchable to accommodate the inflation of the chambers (24).

Abstract: A low cost protective cap that will offer protection of deployable fins for full bore artillery projectiles. A form fitting, composite cap that disintegrates on muzzle exit acts as a protective measure. The cap interfaces with the projectile obturator to form a temporary seal against hot gun gasses. The composition and thickness of the cap allows gas infiltration into the cap material and essentially instant cap degradation upon shot exit.

Abstract: Neutron flux is measured using a subset of detectors (8) selected from a set of detectors (8) introduced into fuel assemblies (2) of the core (1) to obtain measurement signals, the number n of detectors (8) of the subset being at most equal to 15% of the number of fuel assemblies (2) in the core (1), the measurement signals are processed and the instantaneous distribution of the neutron flux in the core (1) is calculated allowing for the measurement signals. At least one operating parameter of the core is calculated from the instantaneous neutron flux distribution, and an alarm is triggered if at least one parameter is outside a fixed range. The operation of the core of the nuclear reactor can therefore be monitored virtually in real time. The neutron flux detectors (8) introduced into the fuel assemblies of the core preferably include measuring probes consisting of collectrons.

Abstract: A skid landing gear for a helicopter, in which the directional stiffnesses of the cross members of the skid landing gear have been de-coupled from one another, such that optimization of the longitudinal stiffniess of the cross members may be performed independently of the optimization of the vertical stiffniess and fatigue life of the cross members. In order to de-couple the stiffnesses in the skid type landing gear, two approaches are employed. In the first approach, the skid landing gear has non-symmetric-section cross members and/or distribution of different materials within the cross-section for de-coupling the vertical stiffness of the cross members from the longitudinal stiffness, such that placement of the ground resonance frequency may be optimized, while retaining the vertical stiffness properties essential for optimizing vertical energy attenuation and fatigue life.

Abstract: A method for providing a communication link from an aircraft (110) to a target (130) is provided. The method includes releasing a first end of a fiber optic cable (120) from an aircraft (110) and then lowering the first end of the fiber optic cable (120) to the target (130). The method also includes connecting the first end of the fiber optic cable (120) to a device at the target (130). The device at the target (130) may then communicate with devices on the aircraft (110).

Abstract: An inflatable restraint system for holding and securing an aircrew member in his seat during ejection from an aircraft has one or more members forming a shield around him or her. The system is controlled to inflate after the aircrew member has started the sequence of ejection and before the rockets that power the seat from the aircraft are ignited. The forces of windblast are deflected by the inflatable restraint from the aircrew member as the aircrew member enters into and remains in the windblast and in his seat.

Abstract: An arrangement in a missile and a method for guiding the missile towards a moving target, for example, an aircraft, wherein the missile has information on its own position, velocity vector, and future velocity characteristic. The missile continuously receives information on the position and velocity vector of the target. The arrangement and method operate such that, a point of interception, at which the missile is expected to strike the target, is calculated from the information on the missile and the target, and a velocity vector of the missile is directed towards the predicted point of interception, which is continuously predicted on the basis of an assumption as to the future movement of the target, and a calculation of the time to impact.

Abstract: A standoff arm and probe assembly for a helicopter has a standoff mounting arm that extends outwardly from the helicopter body in a selected direction, preferably forwardly. The standoff arm has an outer end that mounts a low lateral speed sensing probe that is positioned with an axis substantially parallel to the axis of a helicopter rotor, and within the downwash region of air movement causes by the rotor. The low lateral speed sensing probe has ports that are arranged annularly around the probe, and the pressures sensed at selected annular or peripheral locations on the probe are measured to determine low air speeds. Additionally, the standoff arm can mount a plurality of other sensors such as a pitot pressure probe extending forwardly from the arm, a lateral pressure sensing probe along a lateral side of the arm, and the arm can mount outside air temperature sensors, ice detectors, and total air temperature sensors, if desired.