49results about "Spacecraft vibration damping" patented technology

An apparatus comprises a core having a first surface configured for attachment to a surface of a structure, a face sheet located over a second surface of the core, a number of cavities within an interior of the core, and a number of ports for the number of cavities. The number of ports provides communication between the number of cavities within the interior of the core and the exterior of the core. The number of cavities and the number of ports are configured to reduce noise traveling through the core.

A method and apparatus for deploying a space structure. The space structure is secured to a base with release mechanisms that engage features of the space structure when the base is in a first position. The base is moved from the first position to a second position to deploy the space structure. The release mechanisms are moved at substantially a same time to disengage from the features and release the space structure from the base without imparting a transverse load to the space structure when the base moves from the first position towards the second position to deploy the space structure.

A method and system for calculating a control function for a structural system (10) that can be used to determine an appropriate control force to apply to an active member (18) within a stationary member (12) on the structural system (10). An active member (18) and a stationary member (12) are defined as a two-mass system in which the active member (18) and the stationary member (12) move in opposite directions. The stationary member (12) is mounted to an isolation subsystem (14) that is composed of six isolators (28) at multiple degrees of freedom. The isolation subsystem (14) is softer than the stationary member (12), active member (18) and a spacecraft surface (16) due to a damping element (32) of the isolation subsystem (16). The isolation subsystem (16) is mounted to the spacecraft (16) and decouples the spacecraft (16) from the stationary member (12) and thus the active member (18). An accurate control force for the active member (18) can be determined based upon the above structure (10).

Light-weight acoustic dampening launch vehicle fairing structures having high strength and stiffness and capable of filling to provide high mass density and rapidly unloaded at a predetermined time following launch, and methods of making them are provided. The panels include a monolithic body comprised of fibrous reinforcement material in a polymeric binder and comprise longitudinal chambers extending from the length of body. Fluid passage means are provided at both the top and bottom of the monolithic body and interconnected by the longitudinal chambers for filling and emptying said chambers with a fluid.

A compartment for a space vehicle includes a pressurized structure having a structural wall, the structural wall having interior surfaces facing an interior of the compartment and exterior surfaces exposed to an external environment. An internal mounting structure for mounting a component is provided within the compartment, and mounting features support the internal mounting structure from the pressurized structure. The internal mounting structure is spaced away from the interior surfaces of the pressurized structure, and a thermal fluid is provided in the pressurized structure. The thermal fluid enables convective heat transfer between the component mounted on the internal mounting structure and the interior surfaces of the pressurized structure.

Embodiments of a low frequency isolation device are provided, as are embodiments of a spacecraft isolation system including a plurality of low frequency isolation devices. In one embodiment, the low frequency isolation device includes a three parameter isolator and a break frequency-reducing series spring mechanically coupled in series with the three parameter isolator and having a predetermined axial stiffness (K_{S AXIAL}) and a predetermined lateral stiffness (K_{S LATERAL}). The predetermined axial stiffness (K_{S AXIAL}) of the break frequency-reducing series spring is less than the predetermined lateral stiffness (K_{S LATERAL}) thereof.

The invention discloses a method for inhibiting vibration of satellite solar panel substrates by the aid of solid micro-thruster arrays. The method has the advantages that flexible satellite attitudekinetic equations with the solid micro-thruster arrays are created, joint active vibration controllers for the solid micro-thruster arrays and attitude flywheels are designed, accordingly, the vibration of the satellite solar panel substrates can be quickly inhibited, flexible satellites can be quickly stabilized after being disturbed, and the method is speedy.

An acoustic protection arrangement comprises acoustic protection modules, in particular acoustic absorber modules, wherein at least one of the acoustic protection modules is tuned to a different frequency, in particular has a different peak absorption frequency or acoustic frequency response, than at least another one of the acoustic protection modules. Furthermore, a kit of parts for assembling such an acoustic protection arrangement is provided, and a method for manufacturing such an acoustic protection arrangement is proposed. The acoustic protection arrangement can be implemented based on the kit of parts by selecting appropriate acoustic protection modules and combining them to form an acoustic protection arrangement that fulfils the specific acoustic and geometric requirements for an application, such as a launch vehicle type and/or launch mission. The acoustic protection arrangement can thus be adapted, in particular by tuning the proposed acoustic protection arrangement to achieve a desired overall level of acoustic attenuation.

The attenuation device of the present invention comprises a surface of revolution (11) of circular cross section, such that a set of slots (14) is distributed between the sides of the surface of revolution (11) and adapted to confine an elastic material (15) within the limits defined by the slots (14), generating a structural labyrinth for load passage from a lower interface of the surface of revolution (11) to an upper interface.

The invention discloses an optical effective load passive vibration isolation system, and belongs to the technical field of spacecraft micro-vibration control. The problem that components of an existing vibration isolation platform are prone to be damaged under the condition that the ascending acceleration of a satellite is too high is solved. According to the system, the configuration is characterized in that three upper mounting supports are distributed and fixed to the side face of an upper platform, the positions of the three upper mounting supports are on the same circle, and the centralangle formed by every two upper mounting supports is 120 degrees; six eddy current damper mounting bases are fixed to a lower platform; six supporting legs are divided into three groups, the top endsof the two supporting legs in one group are fixed on the same upper mounting support, one eddy current damper is arranged at the bottom end of each supporting leg, and the eddy current dampers are fixed in eddy current damper mounting bases on the lower platform; and the plane where the two supporting legs in one group are located are perpendicular to the plane where the upper platform is locatedand the plane where the lower platform is located at the same time, the eddy current damper mounting bases at the bottoms of the three groups of supporting legs are distributed on the three edges of aregular triangle, and the eddy current damper mounting bases at the bottoms of one group of supporting legs are located on the same edge.