45results about "Spacecraft coatings" patented technology

A kit for manufacturing a structure in space is provided. In one aspect, the kit includes an inflatable bladder having an internal wall configured to define an internal space of the structure when the bladder is inflated. The kit also includes a structural liner coupled to the internal wall of the inflatable bladder. The kit also includes one or more e-beam physical vapor deposition systems configured to be positioned within the internal space of the structure when the bladder is inflated. The one or more e-beam physical vapor deposition systems configured to form a metallic structural shell on the structural liner.

Disclosed is a satellite which includes an Attitude Determination and Control System (ADCS) and at least one radiative surface, where the orientation of the satellite with respect to the sun is change

A maskless touch up tool having a faceplate and handle. The faceplate is placed on a surface for touching up of a coating on the surface. The surface may be on a rocket body. The faceplate may form a seal around a target region of the surface. The faceplate has an opening configured to receive a touch up coating therethrough for application of the touch up coating to the surface. The handle may extend away from the faceplate and be configured for gripping by a user. The tool may include a reservoir, or an attachment for an external container, configured to collect excess touch up coating from the surface. The tool may be used to touch up the surface coating without the need for masking of the surface.

A space vehicle includes a space vehicle body and a radiator, wherein the space vehicle body includes an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein the space vehicle body includes a panel including a sheet metal; wherein the radiator is positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations, wherein the corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of: sunlight illuminating Earth in a manner of an equinox date, the space vehicle positioned at a low Earth orbit altitude in an equatorial plane, and an observer positioned on the equator, the plurality of corrugations reflect sufficient sunlight away from the Earth such that a minimum magnitude of the space vehicle is in a range of from 6.5 to 9.5.

A space vehicle includes a space vehicle body and a radiator, the space vehicle body including an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, wherein at least a portion of the space vehicle body includes a panel including a sheet metal; the radiator being positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations configured such that, when the space vehicle is evaluated under certain conditions, the corrugations reflect sufficient sunlight away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0.1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle lacking the plurality of corrugations and otherwise equivalent to the space vehicle, when evaluated under the same conditions

The present application provides a thermal control-wave absorbing integrated composite material, a preparation method thereof, and a spacecraft, which relates to the field of aerospace materials technology. The thermal control-wave absorbing integrated composite material comprises a composite ceramic core and a porous coating layer; the porous coating layer covers the composite ceramic core; the chemical formula of the composite ceramic core is La 1‑x Sr x MnO3 / Al2O3, 0.1≤x≤0.3; the porous coating layer includes Al2O3. By introducing Al2O3 into the composite ceramic core, the thermal radiation phase transition temperature can be controlled. The Al2O3 coating layer acts as a shell component to ensure the strength of the thermal control and wave absorbing integrated composite material. The porous structure of the porous coating layer allows radar waves to be absorbed through multiple reflections within the pores, effectively increasing the thermal control and wave absorbing integrated composite material's radar absorption.

To provide a self-supporting space environment protection device, and a method of manufacturing the device.SOLUTION: The device includes a self-supporting polymer-based body manufactured for installation on a component of a space-based system and composed of a space radiation resistant polymer-based material. The device further includes a low-absorptivity finish applied to an external surface of the self-supporting polymer-based body, the low-absorptivity finish having a solar absorptivity (α) value lower than the space radiation resistant polymer-based material, and a low-emissivity finish applied to an internal surface of the self-supporting polymer-based body, the low-emissivity finish having an infrared emissivity (ε) value lower than the space radiation resistant polymer-based material.SELECTED DRAWING: Figure 1

Carbon nanostructure-based composites and methods for making and using the same are described. The carbon nanostructure-based composites can be single-layer or multi-layer composites. Such composites can be useful for shielding radiation, such as that experienced by spacecraft and satellites. Carbon nanostructure-based composites containing low-Z and high-shielding efficiency (e.g., high-Z) materials and having radiation-shielding properties are described herein. A non-limiting exemplary composite is shown in Figure 1.

A self-supporting space environment protection device and method of manufacturing the device are provided. The device includes a self-supporting polymer-based body manufactured for installation on a component of a space-based system and composed of a space radiation resistant polymer-based material. The device further includes a low-absorptivity finish applied to an external surface of the self-supporting polymer-based body, the low-absorptivity finish having a solar absorptivity (α) value lower than the space radiation resistant polymer-based material, and a low-emissivity finish applied to the internal surface of the self-supporting polymer-based body, the low-emissivity finish having an infrared emissivity (ε) value lower than the space radiation resistant polymer-based material.

To provide a system capable of eliminating a complicated function or a high-precision control necessary when a service satellite approaches an outer space non-cooperative object or the like for providing an on-orbit service or removing debris. A cooperative objectification system 1, which cooperatively objectifies an outer space non-cooperative object C existing in an outer space, comprises: a spacecraft S configured to move in the outer space; and a cooperative interface M attached to a surface of the outer space non-cooperative object C by the spacecraft S.

A space vehicle includes a space vehicle body and a radiator, wherein the space vehicle body includes an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein the space vehicle body includes a panel including a sheet metal; wherein the radiator is positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations, wherein the corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of: sunlight illuminating Earth in a manner of an equinox date, the space vehicle positioned at a low Earth orbit altitude in an equatorial plane, and an observer positioned on the equator, the plurality of corrugations reflect sufficient sunlight away from the Earth such that a minimum magnitude of the space vehicle is in a range of from 6.5 to 9.5.

A radiation shielding material for use in spacecraft comprises boron carbide particles dispersed in an epoxy matrix. In one aspect, the material is formed into modular panels attachable to satellite structures for attenuation of neutron and charged particle radiation. In another aspect, the boron carbide epoxy mixture is applied as a conformal coating to satellite components and cured in place to provide localized shielding. The composite may also be used to pot electronic components within a protective volume. The method of fabrication includes mixing boron carbide with epoxy resin and hardener, stirring until homogeneous, and curing the mixture to form a solid shielding structure. The system provides effective radiation mitigation with reduced weight and form factor compared to traditional metal or high-density polymer shields, and is compatible with aerospace integration and environmental requirements.

The present invention relates to a satellite comprising a satellite body including at least one radiation surface configured to radiate heat of the satellite into outer space, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite orbiting around the Earth, wherein the ADCS system is configured to be able to change the orientation of the satellite with respect to the sun between a first position where the radiation surface is away from the sun and a second position where the radiation surface is exposed to the sun. The present invention further relates to a method for controlling heat radiation heat transfer of an artificial satellite.

Provided is a system configured so that when, for example, a service satellite approaches a non-cooperative outer-space object in order to provide an on-trajectory service or to remove debris, it is possible to obviate normally-necessary complex functions or high-precision control. A system 1 for making an object cooperative makes cooperative a non-cooperative outer-space object C which is present in the space of outer space, and the system comprises: an outer-space flying body S which is configured to move in the space of outer space; and a cooperative interface M which is attached to the surface of the non-cooperative outer-space object C by the outer-space flying body S.

Composite materials based on carbon nanostructures and methods of making and using the same are described. The carbon nanostructure-based composite material may be a single-layer or multi-layer composite material. Such composites may be used for radiation shielding, such as those experienced by spacecrafts and space satellites.

The invention discloses a hybrid fiber co-cured multifunctional composite airfoil structure, and belongs to the technical field of aerospace materials. In order to solve the problem that a traditional composite airfoil cannot have the bearing function and the detection function at the same time, a hybrid fiber co-curing process is mainly adopted, an insertion laying layer structure of same-layer butt joint and staggered-layer lap joint is combined, a hybrid fiber transition area is formed, the connection mode is optimized, interference of metal connection on the electrical performance of a functional area is avoided, and the service life of the composite airfoil is prolonged. And meanwhile, the overall rigidity and strength of the airfoil are improved. A traditional composite material airfoil structure is combined with the antenna to form an intelligent skin structure, a functional area is made of a wave-transparent fiber material and is conformal with the antenna, a bearing area is made of a carbon fiber material, and function partition cooperative work is achieved; the hybrid fiber co-curing process is adopted, so that the influence of metallization on the electrical performance of the antenna is eliminated, and the transformation from a traditional composite material structural part to a structural function integrated part is realized.

To provide a system capable of eliminating a complicated function or a high-precision control necessary when a service satellite approaches space debris or the like for removing the debris. A system 1, which removes a space debris C existing in an outer space, comprises: a first spacecraft S and a second spacecraft R configured to move in the outer space; and a cooperative interface M attached to a surface of the space debris C by the first spacecraft S. The second spacecraft R is configured to approach the space debris C using the cooperative interface M, capture the space debris C, and enter the atmosphere together with the space debris C or release the space debris C at a low altitude.

A maskless touch up tool having a faceplate and handle. The faceplate is placed on a surface for touching up of a coating on the surface. The surface may be on a rocket body. The faceplate may form a seal around a target region of the surface. The faceplate has an opening configured to receive a touch up coating therethrough for application of the touch up coating to the surface. The handle may extend away from the faceplate and be configured for gripping by a user. The tool may include a reservoir, or an attachment for an external container, configured to collect excess touch up coating from the surface. The tool may be used to touch up the surface coating without the need for masking of the surface.

A space vehicle includes a space vehicle body and a radiator, the space vehicle body including an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, wherein at least a portion of the space vehicle body includes a panel including a sheet metal; the radiator being positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations configured such that, when the space vehicle is evaluated under certain conditions, the corrugations reflect sufficient sunlight away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0.1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle lacking the plurality of corrugations and otherwise equivalent to the space vehicle, when evaluated under the same conditions

The invention relates to a satellite comprising: a satellite body including at least one radiative surface configured to emit satellite heat to space; and an attitude determination and control ‘ADCS’ system for controlling an orientation of the satellite travelling in orbit around Earth, wherein the ADCS system is configured to change the orientation of the satellite with respect to the sun between a first position in which the radiative surface points away from the sun and a second position in which the radiative surface is exposed to the sun. The invention further relates to a method of controlling thermal radiation heat transfer of a satellite.

Photovoltaic space module (1) with a) a photovoltaic element (3) and b) a spacecraft membrane (2) with a substrate (5) made of a fluoropolymer and a silicon oxide layer (6; 7), wherein c) the photovoltaic element (3) is glued to the spacecraft membrane (2) and d) a cross-section of the spacecraft membrane (2) da) the silicon oxide layer (6) to which the photovoltaic element (3) is adhered, db) the substrate (5) made of the fluoropolymer, dc) a further silicon oxide layer (7) and dd) has an ITO layer (9) or a titanium oxide layer (19), e) wherein a silicon oxide layer (6, 7) is bonded to each of the two sides of the substrate (5), and the silicon oxide layer (6), which is arranged on the side of the substrate (5) facing the photovoltaic element (3), forms an adhesive interface (8) which enables or improves the adhesion of an adhesive layer (4) by means of which the photovoltaic element (3) is bonded to the spacecraft membrane (2), and an ITO layer (9) is bonded to the outside of the silicon oxide layer (7) on the side of the substrate (5) facing away from the photovoltaic element (3).

The present invention relates to a satellite comprising a satellite body including at least one radiation surface configured to radiate heat of the satellite into outer space, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite orbiting around the Earth, wherein the ADCS system is configured to be able to change the orientation of the satellite with respect to the sun between a first position where the radiation surface is away from the sun and a second position where the radiation surface is exposed to the sun. The present invention further relates to a method for controlling heat radiation heat transfer of an artificial satellite.

Techniques for providing a heatshield involve receiving a starting portion of a tape from a tape supply, positioning the starting portion of the tape in contact with a tool structure, and after the starting portion of the tape is positioned in contact with the tool structure, and robotically moving the tool structure and a tape deployment head relative to each other to precisely guide the tape onto the tool structure. Robotically moving the tool structure and the tape deployment head relative to each other includes receiving a set of sensing signals indicating current position of the tool structure and the tape deployment head relative to each other, and based on the set of sensing signals, applying the tape under pressure (e.g., via pressure and temperature control) to form the heatshield.

Carbon nanostructure-based composites and methods of making and using thereof are described. The carbon nanostructure-based composites may be single-layered or multi-tiered composites. Such composites can be useful for radiation shielding, such as experienced by spacecraft and space satellites.

A maskless touch up tool having a faceplate and handle. The faceplate is placed on a surface for touching up of a coating on the surface. The surface may be on a rocket body. The faceplate may form a seal around a target region of the surface. The faceplate has an opening configured to receive a touch up coating therethrough for application of the touch up coating to the surface. The handle may extend away from the faceplate and be configured for gripping by a user. The tool may include a reservoir, or an attachment for an external container, configured to collect excess touch up coating from the surface. The tool may be used to touch up the surface coating without the need for masking of the surface.

The present invention discloses a rapid prototyping heat-resistant film and a preparation method thereof. The steps include: uniformly dispersing a reinforcing filler and ground fibers in silicone rubber to obtain a mixture; performing vacuum kneading on the mixture; and then calendering the kneaded and dispersed mixture to obtain a preliminary formed film. Finally, the preliminary formed film is cured to obtain a heat-resistant film. The heat-resistant film obtained by the present invention can replace the traditional air-sprayed silicone rubber heat-resistant coating construction, and is particularly suitable for application to aircraft surfaces to achieve the purpose of protecting against aerodynamic heat. The obtained structural parts have good mechanical, thermophysical properties, and ablation resistance. At the same time, compared with the air-sprayed construction of traditional heat-resistant coatings, the present invention can avoid the use of a large amount of solvents and significantly shorten the construction period. This has far-reaching significance for the rapid delivery of future heat-resistant products and the promotion of environmental protection.

The present invention is directed to a phosphated epoxy resin comprising at least one terminal group comprising a phosphorous atom covalently bonded to the resin by a carbon-phosphorous bond or by a phosphoester linkage; and at least one carbamate functional group. The present invention is also directed towards aqueous resinous dispersions comprising the phosphated epoxy resin, methods of coating a substrate, coated substrates, and methods of making a phosphated epoxy resin.

To provide a system capable of eliminating a complicated function or a high-precision control necessary when a service satellite approaches space debris or the like for removing the debris. A system 1, which removes a space debris C existing in an outer space, comprises: a first spacecraft S and a second spacecraft R configured to move in the outer space; and a cooperative interface M attached to a surface of the space debris C by the first spacecraft S. The second spacecraft R is configured to approach the space debris C using the cooperative interface M, capture the space debris C, and enter the atmosphere together with the space debris C or release the space debris C at a low altitude.

Coatings and materials that are atomic oxygen resistant and have an atomically smooth surface that can reduce drag are disclosed. The coatings and materials can be used on at least a portion of a spacecraft intended to operate in harsh environments, such as stable Earth orbits at about 100 km to about 350 km.