77results about "Protection against meteorites" patented technology

An aerospace hardware derelict item is salvaged in space, modified, and reused to provide manned facilities in orbit. The hardware packages added onto the salvaged discarded item enhances its value, and incorporates at least the subsystems required to effectively reuse portions of previously discarded launch vehicle components, and other derelict objects in space. The hardware, and technique used reduces the cost of launching comparable hardware to orbit, because of the reuse, and provides a human habitation in orbit. The salvaged items include the external tank of the space shuttle, other derelict orbital hardware, the add-on cargo pod in two forms, and the ability to convert the derelict into a cost effective reusable item. The salvaged hardware is initially capable of contributing mass, length, interior volume, strongback, rotational stability mass, interior pressurized volumes, artificial gravity, and stability with the addition of simple subsystems for salvage, and interior development packages.

Fiber-based debris interceptors are used to intercept and / or contain space debris. The debris interceptors may be made up of fibers that are formed in space from a material supply on a space vehicle. The fibers for the debris interceptors may be formed by extrusion, with a heat source, such as a mirror to focus sunlight, used to heat the material of the material supply that is carried on the space vehicle. The debris interceptor may be separated from the space vehicle, and used to remove debris from an orbit, or otherwise prevent debris entering an orbit from damaging a satellite or other spacecraft that travels in that orbit. The debris interceptor may be deployed prior to later launch of a valuable spacecraft, in order to “cleanse” the intended orbit of debris. Debris objects may pass through the debris interceptor, but in so doing may lose energy so as to de-orbit.

A multi-layer cavity-type aeration for expansion protection mechanism relates to an aeration for expansion protection mechanism, which aims at solving the problems that a monolayer protection screen can not effectively avoid the damage caused by the collision from the space chippy fraction, and a mechanism-typed expansion protection mechanism has low folding efficiency, low expansion reliability and a small volume after expansion. Parts of the periphery of the adjacent two protection screens cling together to form a cavity-type aeration for expansion protection screen to be rigidified, which is connected with an aeration supporting frame to be rigidified; a main inflation pipe is arranged in the cavity-type aeration for expansion protection screen to be rigidified and communicated with an air-discharging pipe; the lower end of the aeration supporting frame to be rigidified is provided with an aeration supporting pipe to be rigidified, which is communicated with the inflation pipe of a supporting pipe; two ends of an aeration bent pipe are communicated with an aeration pipe to be rigidified and the aeration supporting pipe to be rigidified. The invention can avoid the damage caused by the collision from the space chippy fraction, and has the advantages of high folding efficiency, high expansion reliability and a large volume after expansion.

Provided herein methods and structures for construction of light weight sandwich panels or load-bearing panels having improved ballistic protection. The methods disclosed can facilitate construction of a hybrid core comprising ballistic fabrics and deformable pins that exhibit superior resistance to ballistic penetration. The methods and structures of the present disclosures can be advantageously used in many applications, such as armored vehicles.

A spacecraft for removing space debris is disclosed. The spacecraft includes a satellite bus, a shield member foldable on an outer side face of the satellite bus and disposed facing towards space debris to reduce a movement speed of the space debris, and a support member configured to support the shield member with respect to the satellite bus, in which the shield member includes a central panel configured to overlap one face of the satellite bus, a plurality of first panels connected to peripheral sides of the central panel and radially extended, and a plurality of second panels located between the first panels.

A defense mechanism which can be charged and unfolded on track and is capable of resisting the impact from space debris and micrometeoroid relates to a defense mechanism for resisting the impact from space debris and micrometeoroid, aiming at solving the problems of high cost and big difficulty in large size defense screens assembled in space for one or more launches and the problems of low folding efficiency, low reliability and small size after being unfolded on mechanically unfolded defense screens. Two adjacent layers of plane frames which can be hardened and can be unfolded by charging are connected through a support upright pipe assembly which can be unfolded by charging, and defense screens are arranged in the plane frames which can be hardened and can be unfolded by charging and are connected with the plane frames which can be hardened and can be unfolded by charging; horizontal leads and air inlet pipes are arranged inside the plane frames which can be hardened and can be unfolded by charging; vertical parallel leads and charging pipes are arranged inside the support upright pipe assembly which can be unfolded by charging; the horizontal leads are connected with the vertical parallel leads; the first charging pipe and the second charging pipe are communicated with short and long support pipes which can be hardened and can be unfolded by charging; the vertical charging pipes are communicated with the support upright pipes which can be hardened and can be unfolded by charging. The defense mechanism is good in defending effect, high in reliability, light in weight, high in folding efficiency and low in cost.

An apparatus having a composite sea-based structure with a first carbon nanotube infused material and a second carbon nanotube infused material. The first and second carbon nanotube infused materials each having a range of carbon nanotube loading selected to provide different functionalities.

A retractable rigid dome for protecting a feature, such as a docking mechanism, a hatch or other equipment on an exterior surface of a space vehicle, includes a plurality of rigid arcuate segments, each having opposite ends respectively pinioned at opposite sides of the feature at about the surface of the vehicle for rotational movement about an axis of rotation extending through the opposite ends thereof and through an arcuate path of rotation extending over the feature. The radial sizes of the segments are staggered such that, in a fully deployed position of the dome, in which adjacent segments are rotated apart from each other at a maximum relative angle therebetween, the segments combine to form an arcuate shield over the feature, and in a retracted position of the dome, in which adjacent segments are rotated together at a common angle relative to the surface of the vehicle, the segments are nested in radial alignment with each other.

An apparatus having a composite air-based structure with a first carbon nanotube infused material and a second carbon nanotube infused material. The first and second carbon nanotube infused materials each having a range of carbon nanotube loading selected to provide different functionalities.

A device for trapping space debris with at least one rope-like trap for space debris is provided. The trap includes an end piece configured for being wound around space debris or parts thereof. The rope-like trap is further configured in such a manner that upon contact with the space debris, it can generate a friction that is sufficient for towing the space debris.

The invention relates to a protection system that resists space junk and micrometeoroid, rigidifies and space inflates in orbit, which aims at solving the problems of high cost and sheer difficulty for big-size protection screen space-assembly for single or multiple launching, the problems of low folding efficiency, low reliability, small unfolded size in mechanical protection screen unfolding, and the problems of large folding size, inflexible folding, large launching size and small unfolded size of rigid protection screens. Two adjacent layers of rigid space inflating plane frame are connected by the rigid space inflating support vertical tube assembly; protection screens waiting for rigidifying are arranged in the rigid space inflating plane frame; a first and a second inflating pipe are arranged in the rigid space inflating plane frame; vertical inflating pipes are arranged in the rigid space inflating support vertical tube assembly. The protection system has the advantages of simple structure, low cost, high reliability, light weight, high folding efficiency, small folding size, flexible folding and small launching size, which can applied to aircrafts such as satellites, spacecrafts and space stations.

Methods and apparatus are provided for a shield to protect a surface from the impact of hyper-velocity projectiles. The apparatus comprises a foam material that is configured to cover the surface to be protected and is attached directly to that surface. A coating material is typically disposed on the outer surface of the foam material and may penetrate the foam material to a predetermined depth. The foam material and the coating material are selected to form a composite having predetermined values of sonic velocity, toughness, and thermal conductivity. The composite of foam material and coating material can be significantly lighter in weight than a metal shield having equivalent protective characteristics.

A folding, retractable dome for protecting a feature, such as a docking mechanism, a hatch or other equipment at an exterior surface of a space vehicle, includes a plurality of arcuate ribs, each having opposite ends respectively pinioned at opposite sides of the feature at the surface of the vehicle for rotational movement about an axis of rotation extending through the opposite ends and through an arcuate path of revolution extending over the feature, and a flexible cover attached to each of the ribs such that, in a deployed configuration of the dome, in which adjacent ribs are rotated apart from each other at a maximum relative angle therebetween, the cover is stretched generally tangentially between the adjacent ribs to form a generally arcuate shield over the feature, and in a retracted position of the dome, in which adjacent ribs are rotated together at a minimum relative angle therebetween, the cover is collapsed to define folded pleats between the adjacent ribs.

The invention discloses a monitoring method and a system for an on-orbit spacecraft suffering from space debris collision. The monitoring method comprises the following steps that corresponding target ultrasonic signals are obtained through a plurality of ultrasonic sensors, and then the corresponding target ultrasonic signals are subjected to amplification and gain treatment and ultrasonic monitoring data is obtained; when any of the ultrasonic monitoring data is judged to exceed a presupposition trigger threshold, specific ultrasonic monitoring data is gathered and stored; the specific ultrasonic monitoring data is the ultrasonic monitoring data of all the ultrasonic sensors in the presupposition trigger duration; when all the ultrasonic monitoring data is judged not to reach the presupposition trigger threshold, the stored specific ultrasonic monitoring data is read; and analysis and calculation are conducted according to the specific ultrasonic monitoring data, then the collision position and the collision strength of the on-orbit spacecraft are obtained. According to the method and the system for the on-orbit spacecraft suffering from the space debris collision, acoustic emission monitoring and location produced by a space debris collision structure are achieved through a passive ultrasonic monitoring form.

A shear-thickening fluid is used in conjunction with fabrics utilized in an expandable spacecraft. The combination of the fluid and the fabric allows the fabric to resist penetration by hypervelocity particles in space.

Various embodiments provide multi-layered self-healing materials, capable of repairing puncture damage. The multi-layered self-healing materials, capable of repairing puncture damage of the various embodiments may be constructed by sandwiching a reactive (e.g., oxygen sensitive) liquid monomer formulation between two solid polymer panels, such as a polymer panel of Barex 210 IN (PBG) serving as the front layer panel and a polymer panel of Surlyn® 8940 serving as the back layer panel. The various embodiments may provide methods to produce multi-layered healing polymer systems. The various embodiments may provide a two-tier, self-healing material system that provides a non-intrusive capability to mitigate mid to high velocity impact damage in structures.

Space traffic is managed using data gathered by orbital sensors. A constellation of near-equatorial orbiting satellites can be established, with each satellite in the constellation including at least one sensor for tracking resident space objects (“RSO”). The tracking data gathered by these orbital sensors can be fused with previously-gathered orbital tracking data and / or tracking data from ground-based sensors and used to adjust orbital information for the RSOs. The adjusted orbital information for the RSOs can, in turn, be used to issue conjunction warnings, to adjust the orbits of one or more satellites in the constellation (e.g., to intercept a debris object; to intercept a target; to avoid an active spacecraft), and / or to adjust the orbits of one or more other spacecraft (e.g., to avoid debris).

An impact resistant insulation tile is provided that is capable of withstanding the impact of Micrometeriods and Other Debris (MMOD). The tiles are secured to an exterior surface of a Reusable Launch Vehicle (RLV). A durable coating is applied to the exterior surface of a thermally insulative layer to fragment and slow down MMOD that collide with the tile. In addition, a shock layer may be embedded within the thermal insulator to further reduce the size and speed of MMOD particles. A ballistic strain isolator pad may also be located between the thermal insulator and the RLV structure to capture fragmented particles.

An inflatable airlock for use with a spacecraft is disclosed. The airlock has a substantially cylindrical shaped layered shell comprised of an outer meteoroid shield layer, a restraint layer under the meteoroid shield, and an air barrier layer under the restraint layer. There is a door arrangement attached to the airlock and the airlock is adapted to being attached to a spacecraft bulkhead that also includes a door. When inflated, a person can open the spacecraft bulkhead door and pass through from the spacecraft into the airlock or vice versa. When the atmosphere is removed from the airlock, a person can pass from the airlock into space or vice versa.

A protective assembly comprises a first region formulated and configured to provide protection from alpha, beta, and electromagnetic radiation and comprising a composite of particles and polymer; a second region formulated and configured to provide protection from ballistic impact and comprising a composite of fibers and polymer; and a third region formulated and configured to provide protection from thermal radiation and comprising a composite of particles, fiber, and polymer. The protective assembly may be provided on an aerospace structure. The protective assembly may be formed on the aerospace structure body using a co-curing process.

A solar generator for spacecraft or satellites, including a solar panel, a detector layer, a first evaluation means and a second evaluation means. The solar panel includes a plurality of solar cells. The detector layer includes first conductors substantially in parallel in a first plane and second conductors substantially in parallel in a second plane. The second plane is substantially in parallel to the first plane such that the first conductors are at an angle to the second conductors. The first evaluation means is for detecting a failure of one or more of the plurality of solar cells. The second evaluation means is for testing electrical status of the first conductors and the second conductors to determine whether the failure detected by the first evaluation means was caused by a space object damaging or severing one or more of the first conductors and the second conductors.

One aspect of the present invention is a puncture healing polymer blend comprising a self-healing first polymer material having sufficient melt elasticity to snap back and close a hole formed by a projectile passing through the material at a velocity sufficient to produce a local melt state in the first polymer material. The puncture healing polymer blend further includes a non-self-healing second material that is blended with the first polymer material. The blend of self-healing first polymer material and second material is capable of self-healing, and may have improved material properties relative to known self-healing polymers.

A spacecraft carbon nanotube shield is disclosed. Shield segments are produced in a facility in space. The segments are transported from the facility to a vicinity of a spacecraft hull. The segments are assembled over the hull to substantially cover an area of the hull.

Radiation shields include an outer composite wall, an inner composite wall, and an enclosed fuel chamber disposed between the inner and outer composite walls. A fuel may be contained within the enclosed fuel chamber. Spacecraft include one or more such radiation shields, and may further include a fuel cell and a conduit providing fluid communication between the fuel chamber of the radiation shield and the fuel cell. Methods of forming a radiation shield system include forming an enclosed fuel chamber between an outer composite wall and an inner composite wall of a radiation shield, and providing fuel within the enclosed fuel chamber.