106results about "Cosmonautic component separation" patented technology

This invention provides a satellite-launch separation signal system and a satellite-launch separation system, comprising: a first user unit that outputs two satellite-launch separation signal sources, each leading to two different contacts of a first analog separation relay; these signals then return to the first user unit via a first separation connector and a second separation connector; the input and output terminals of the first analog separation relay are each connected to a satellite display socket via cables; and a second user unit that outputs two satellite-launch separation signal sources, each leading to two different contacts of a second analog separation relay; these signals then return to the second user unit via a first separation connector and a second separation connector; the input and output terminals of the second analog separation relay are each connected to a satellite display socket via cables. This invention ensures that the signal is provided by the analog satellite-launch separation relay during ground testing, and avoids fatal malfunctions such as accidental separation commands issued before launch or mis-separation due to severe vibrations during the active phase, thus improving the reliability of the satellite-launch separation signal link.

A reusable orbital vehicle (14) for a space transport system (10) comprises a fuselage (22) the shape of which widens in a direction extending from the front end (24) of the reusable orbital vehicle to a rear end (26) of said fuselage (22) and inside which there is defined a housing (32), and a crew-evacuation vehicle (34) housed in said housing (32) and oriented in such a way as to be extracted from the housing (32) in a direction (D1) directed from said rear end (26) of the fuselage (22) toward said front end (24) of the reusable orbital vehicle (14).

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

This application relates to the field of aerospace technology, and in particular to an irregularly shaped multi-satellite distributor. The lower end face of its irregularly shaped load-bearing structure is abutted against the upper end face of its conical load-bearing structure. The lower end face of the irregularly shaped load-bearing structure is circular, and the upper end face is elliptical. Its load-bearing beam is fixed within the mounting cavity of the irregularly shaped load-bearing structure. A top satellite mounting bracket is fixed to the upper end of the irregularly shaped load-bearing structure, and a side satellite mounting bracket is fixed to the outer side of the irregularly shaped load-bearing structure, perpendicular to the major axis direction of the upper end face. An irregularly shaped satellite mounting bracket is fixed to the outer side of the irregularly shaped load-bearing structure, perpendicular to the minor axis direction of the upper end face. Regularly shaped satellites are fixedly mounted on the top satellite mounting bracket and / or the side satellite mounting bracket, and irregularly shaped satellites are fixedly mounted on an irregularly shaped satellite mounting bracket. This application allows the protruding points of the satellites to be enveloped within a safe envelope range, and avoids increasing the size of the fairing.

This provides a coupling mechanism that can reliably protect the internal mechanism from adverse effects of the external environment. [Solution] The coupling portion 3A of one robot module 2A comprises an annular base portion 4 having an opening in the center, a planar flexible sealing member 4 covering the opening of the base portion 4, an engaging portion 7 disposed outside the sealing member 4, and a drive unit 9 built inside the base portion 4 separated by the sealing member 4. The coupling portion 3B of the other robot module 2B comprises an engaged portion 7' ​​that engages with the engaging portion 7. The drive unit 9 engages the engaging portion 7 via the sealing member 4, and the sealing member 4 deforms in sync with the movement of the engaging portion 7.

A satellite capture module (1) comprising: a main structure (10) extending around a main axis (X); a deployment system (40) configured to move the main structure (10); a plurality of arms (20) mounted on the main structure (10) around the main axis (X); a plurality of anchoring members (30) configured to engage an anchoring structure (103) of a target satellite (100), each anchoring member (30) being mounted on a respective arm (20) of the plurality of arms (20); wherein each arm (20) is movable between a first configuration, in which the anchoring member (30) is placed at a first radial distance from the main axis (X), and a second configuration, in which the anchoring member (30) is placed at a second radial distance from the main axis (X) greater than the first radial distance.

Disclosed are systems and methods for a distributed space transportation network. Satellite launches to orbit are more efficiently performed by large rockets. Modern satellites are in smaller form factor, leaving the large launch rockets with excess capacity. Small satellite operators can use ride-shares, but do not have efficient options for delivering their satellites to their desired destination and may be forced to operate their satellites in compromise orbits. The disclosed distributed space transportation network maintains a fleet of space tugs, which can dock with satellites in space at an initial arrival destination and transport the satellites to their final destinations. In one embodiment, the space tugs can dock with satellite depots to obtain fuel and repairs.

The invention belongs to the technical field of aerospace controlled separation mechanisms, and relates to a hot knife unlocking device. The heat knife comprises a heat knife base, and a heat knife containing cavity is formed in the heat knife base. When fusing unlocking does not need to be conducted, the adjustable protection structure rotates to the position above the heating rod, the upper portion of the heating rod is covered and protected through the adjustable protection structure, the situation that when an accident occurs, a part collides with the heating rod, and consequently the heating rod cannot conduct heat knife fusing normally is avoided, and safety and stability are guaranteed; the heat knife anti-bounce piece can ensure that the lifting power supply module and the heating rod are prevented from bouncing in a normal state, the lifting power supply module and the heating rod move towards the rope through the elastic connecting assembly, elastic release is achieved, moving elastic impact force and high temperature are carried, heat knife fusing is rapidly carried out on the rope, and the heat knife anti-bounce piece can be repeatedly used; the fusing stability of the hot knife is ensured, the failure rate is greatly reduced, and the practicability is high.

Spacecraft servicing devices and related methods may include a propellant tank configured to store a propellant and to be placed into fluid communication with a portion of the target spacecraft.

This invention discloses a locking and release device for a 4D-printed small satellite based on SMA (Super-Magnetic Atomizing) technology, belonging to the fields of aerospace and biomimetic 4D printing technology. The device consists of a main supporting shell 1, a launch platform 2, an auxiliary drive device 3, a locking lever 4, a locking and release spring assembly 5, and an energy storage spring assembly 6. The auxiliary drive device 3 controls the downward movement of the locking mechanism through temperature changes of the inner and outer springs, thereby compressing the energy storage spring assembly 6 and heating it to achieve autonomous energy storage. The launch platform 2 supports the satellite at its upper end and is tightly connected to the energy storage springs at its lower end. The locking and release spring assembly 5 is used to lock the satellite and drive its release. In summary, this invention provides a locking and release device for a 4D-printed small satellite based on SMA, making the failure modes of the satellite launch device effectively controllable and improving the device's adaptability to the extreme environment of space.

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

A rocket stage having a frame is disclosed. The frame includes a nose part and a mid-section part. The frame is configured to separate into two or more gliding parts after the stage has fulfilled its main purpose. After separation, the gliding parts glide down to the ground. The gliding part comprises a section of the nose part and a section of the mid-section part and a flat gliding surface alongside the section of the mid-section part. The slope of the nose part of the gliding part is steeper than the slope of the mid-section part of the gliding part. The air flows around the frame of the gliding part and a vortex forms above the upper edges of the gliding part, forcing the gliding part in the gliding position when it is in the atmosphere.

This invention provides a large-scale load-bearing mechanical synchronous linkage unlocking mechanism, comprising a base plate, a retainer, a top plate, several connecting rods, several blades, and several leaf springs. The beneficial effects of this invention are: the mechanism has the function of locking a threaded clamping rod; the mechanism has a high load-bearing capacity (not less than 10,000 Newtons); the mechanism has a self-locking safety protection measure in the clamped state; the mechanism has a compact space and a design similar to a camera shutter; the mechanism has the functional feature that the unlocking direction and the unlocking unfolding direction are perpendicular to each other.

“Space launcher equipped with an auxiliary propulsion system and auxiliary propulsion system equipping a space launcher” Space launcher (1) comprising a first stage equipped with a first propulsion system dedicated to the intra-atmospheric flight phase of the launcher and a second stage (E2) designed to receive a payload, comprising a main engine (10) and an auxiliary propulsion system (APS) integrating two auxiliary propulsion units (APUs) (13, 16) and a set of reaction control thrusters (RCS) responsible for attitude control of said second stage during the extra-atmospheric flight phase of the launcher. See Figure 1

To provide a method of autonomous maneuver generation to mate connectors.SOLUTION: A method includes providing an image of feature data on the basis of a feature extraction model of connector mating. The image depicts a first portion of a first device and a second portion of a second device. The feature data includes coordinates representing key points of each of the first and second devices depicted in the image. The method also includes obtaining position data indicating a position in 3D space of a connector of the first device. The method further includes providing the feature data and the position data to a trained autonomous agent to generate a proposed maneuver to mate the connector with a connector of the second device.SELECTED DRAWING: Figure 1

A small satellite docking adapter includes a probe adapter, the probe adapter including: one or more latches; one or more springs; and a cone adapter, the cone adapter including: one or more grooves, wherein: the one or more latches are configured to couple to the one or more grooves; and the one or more springs are configured to cause the one or more latches to engage with the one or more grooves to lock the probe adapter to the cone adapter, and to disengage with the one or more grooves to unlock the probe adapter and the cone adapter.

The application provides a reliable separation system and method for separating a service space load, which comprises an electrical separation connector, a cutter, an anti-falling pyrotechnic lock, a power supply / charge management module and a separation controller; the electrical separation connector is a main means for realizing electrical separation of a satellite platform and a separation load; the cutter is a backup means for realizing electrical separation of the satellite platform and the separation load, and is used for cutting a cable to realize separation when the electrical separation connector cannot normally separate; the anti-falling pyrotechnic lock is used for connecting a plug and a socket of the electrical separation connector; the power supply / charge management module is used for managing charging and power supply of the separation load; and the separation controller is used for controlling the electrical separation connector, the cutter and the anti-falling pyrotechnic lock. The application aims at the problem that the traditional electrical separation connector cannot normally separate due to vacuum cold welding of the separation load in long-term on-orbit, and ensures reliable separation by designing the cutter as a backup means.

Provided is a system, method and device for spacecraft connections. The system includes a first spacecraft including a structural connector that provides structural support and alignment to a second structural connector and an electrical connector that electrically connects the spacecraft platform to a launch vehicle and / or a second spacecraft. The electrical connector includes a first connector pad that mates with a second connector pad. The electrical connector is absent a connection force inherent in the electrical connector, wherein the connection force is a force that keeps the first connector pad in contact with the second connector pad.

A first space vehicle (fig.1,101) has two or more latch arms 151 and a moving probe 163 carried by a first valve 160. The arms are rotated and translated to couple the spacecraft and to draw a second

A frame for a satellite is provided. The frame includes a body and posts that are constructed of aluminum. The body can receive spacecraft components therein. The posts extend from the body and are operable to releasably couple with posts from another frame for another satellite and / or with a payload adaptor fitting.

The application provides a high-compact integrated interface device suitable for on-orbit assembly of a spacecraft, comprising a passive end assembly and an active end assembly, the passive end assembly comprising a passive end shell, a first permanent magnet and a locking piece, and the active end assembly comprising an active end shell, an electromagnet, a second permanent magnet, a locking assembly and an excitation control circuit; the electromagnet is arranged between the first permanent magnet and the second permanent magnet, and the magnetic pole arrangement directions of the first permanent magnet and the second permanent magnet are opposite; the locking piece is mounted on the end face of the passive end shell, and the locking assembly is correspondingly mounted in the active end shell. The application controls active adsorption and active separation between the passive end assembly and the active end assembly by passing positive current or reverse current to the electromagnet, has small space occupation, high compactness, is suitable for on-orbit docking of a small module, needs to consume energy only at the moment of switching between the locking and separation states, does not need to consume energy in normal state, and improves the service life.

Quick disconnect devices for high pressure fluid transfer, and associated systems and methods are disclosed. A representative quick disconnect system includes a first connector and a second connector. The second connector can have an opening sized and shaped to receive a first end of the first connector. The second connector can include a poppet positioned to open the first connector when the first connector is connected to the second connector. The second connector can include an inner sleeve moveable between a first position wherein the poppet head forms a fluid-tight seal with the annular seat of the inner sleeve, and a second position wherein the second end portion is open to permit fluid flow through the end portion of the inner sleeve. In some embodiments, the inner sleeve is pressure balanced in every direction.

The present invention relates to a spacecraft 1 for supplying electric energy to a client ship 2 located in free space, in orbit and / or on a celestial body, the spacecraft 1 comprising a main structure 10 having an electric thruster 10A, a chemical thruster 10B and a solar electric energy generator 10C, a first fuel container 11 for fuel for the electric thruster 10A and a second fuel container 12 for fuel for the chemical thruster 10B, the spacecraft 1 being such that the main structure 10 can alternatively be connected or separated to the first fuel container 11 or the second fuel container 12, the first fuel container 11 and the second fuel container 12 can be connected to or separated from each other, and the solar electric energy generator 10C is intended to be deployed or stored.

The present application relates to the technical field of space robot, in particular to a kind of three degrees of freedom soft docking joint, including transmission mechanism, damping buffer component and sensing component.The transmission mechanism includes clutch, cross shaft structure, Z-axis rotating shell, Y-axis rotating shell etc.;Damping buffer component includes rotary magneto-rheological damper and torsional spring mechanism etc.;Sensing component includes encoder etc.The soft docking joint mainly utilizes damping buffer component to control the flexibility of docking mechanism, according to operation task to realize space six-dimensional momentum unloading, to realize the soft docking of space operation.