RU2514981C2 - Spacecraft undocking system - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to aerospace engineering, particularly, to craft docking and undocking means. Spacecraft undocking system arranged between carrier rocket mainframe comprises housing with locks, pusher and solar battery wing attachment assemblies. Every solar battery wing attachment assembly is arranged at undocking system housing, under solar battery wing, and consists of a support bracket with vertically-adjustable threaded rods secured thereat to make the thrusts. Threaded rod lower ends are secured at said support bracket while their upper ends are connected via spherical tips with pressure plates. Rubber gasket is fitted between pressure plate surface at every said adjustable thrust and bottom end surface of appropriate panel of solar battery wing.

EFFECT: decreased impact and vibratory stresses.

2 cl, 5 dwg

Description

The invention relates to systems for dumping or separating objects, mainly spacecraft (SC) and other payloads (ST) from supporting structures (adapters or instrument compartments) of launch vehicles (LV) when they are put into the calculated orbit, and can be used in the field rocket and space technology.

A known separation system (CO) of the spacecraft according to the patent of the Russian Federation No. 2268208 from 01.20.2006, IPC B64G 1/64, containing a detachable holding device installed between the supporting structure of the pH (for example, an adapter or instrument compartment of the pH) and the spacecraft, consisting of a body , locks and pushers interacting with the instrument cluster of the spacecraft and ensuring its separation from the launch vehicle. In this case, the locks hold the spacecraft on the LV support structure when it is put into operation orbit, and the pushers are designed to separate the spacecraft from the spacecraft. Elements (locks and pushers) of the SD of this design interact only with the instrument panel (PB) of the spacecraft, while the solar battery (SB), consisting of panels rolled up in a package and mounted on one of the sides of the spacecraft, hangs cantilever, loading nodes for fixing the SB to the spacecraft with significant tearing loads. When fastening the SB wing (fastening of the SB wing, as a rule, is carried out to the side of the SC SC), the CO locks (locks of SC fastening to the LV) also have significant loads due to the large eccentricity between the center of mass of the SC wing and the longitudinal axis of the SC SC. Significant loads acting on the CO locks and attachment points of the SB fixation to the spacecraft can adversely affect their performance when the spacecraft is separated from the LV and when the SB wing is deployed to the working position after the spacecraft is separated from the LV. This is a drawback of CO if the SC is fixed on it, which has a multi-link SB wing, consisting of several panels and fixed to the side of the SC SC.

A known satellite separation system and a device for holding SB according to US patent 3327967, cl. 244-1 dated 03/31/1965, in which the SB panel is mounted on the satellite (SC) body and mounted on the LV with attachment (fixation) nodes. The SB panel is in the folded (transport) position and, after separation of the satellite from the LV, is deployed relative to the satellite body to the operating position.

An adapter is also known - a separation system for docking with a LV of an aircraft (LA) according to the patent of the Russian Federation No. 2198117 dated 02/10/2003, IPC B64G 1/00, B64G 1/44. (application No. 99102341/28 of 02/08/1999). The adapter — the separation system consists of a body (in the form of a shell with a stringer-and-ring kit), on which locks (clamps) are installed that connect the aircraft with the LV and ensure their separation in accordance with the flight program. An aircraft includes a housing on which hinged structures are installed, for example, SB panels having attachment points mounted outside the aircraft and located on an adapter (on the CO housing) mounted on the LV. The adapter (CO) is connected by locks (latches) to the aircraft and rigid nodes with LV (after the locks are activated and the aircraft is separated from the LV, the adapter (CO) remains on the LV). In addition, fasteners of the hinged structure (SB panels) are made on the adapter (СО). Those. SB panels have fasteners with both the aircraft body (upper nodes) and the adapter (СО) of the PH (lower nodes). After the head fairing is reset, before the separation of the aircraft from the LV, the fastening nodes of the SB panels on the adapter (lower nodes) are first unlocked, and then the CO locks (adapter clamps) are activated, which ensure separation of the aircraft from the LV. After separation of the aircraft from the LV, the SB panel is held and swiveled relative to the aircraft body at the upper fixation nodes.

The disadvantage of these JI designs that implement the fastening schemes of the SC SC panels on the SC body and on the adapter housing (CO) installed on the LV is that they provide a certain sequence of operation of the SB fastening nodes to the PH adapter and the CO locks of the SC fastening to the PH. In the event of a delay in the operation (or failure) of the attachment points of the SB to the LV adapter, the spacecraft will not separate from the LV or separate with the inevitable damage to the SB panel held by the broken mount. The disadvantage of securing the wing or the SB panel of a large area on top of the spacecraft body (on the spacecraft instrument unit), and on the bottom on the adapter (on the CO) of the LV leads to a large inter-bearing distance and, consequently, to large deflections and movements of the SB wing relative to the spacecraft AS during loading during operation. Fastening the SB panel using locks on the satellite housing and on the adapter housing, i.e. structures with different flexibility can also lead to unacceptable deformations of the SB panel design and the failure of the photoelectric converters installed on it.

In addition, after unlocking the lower attachment points of the SB panel (or wing) to the LV adapter, the remaining attachment only from above to the SC CA does not provide reliable fastening of the panel (wing), therefore, when the SC is separated from the PH, the SB panel (wing) can strike the lower part of the SC KA.

As a result of the analysis of the patent and scientific and technical literature, the technical solution according to the patent of the Russian Federation No. 2198117 dated 02.10.2003, IPC B64G 1/00, B64G 1/44, which describes the SOA having a body with the locks (and pushers) mounted on it to the spacecraft to the launch vehicle and the fastening points of the spacecraft to the body of the aircraft.

The objectives (goals) of the proposed device of the spacecraft separation system are:

- reduction of loads on the fixation nodes of the wing of the SB to the SC CA and the locks of the SC fastening the SC to the LV;

- an increase in the rigidity of the SB wing installation on the SC SC; reduction in deflections and movements of the SC wing relative to the SC SC;

- improving the reliability of the separation of the spacecraft from the launch vehicle.

To achieve the objectives (goals) in the known device of the spacecraft separation system installed between the supporting structure of the spacecraft and spacecraft and containing the SB wing mounts, these SB wing mounts are installed on the body of the SB under the SB wing and consist of support brackets with adjustable brackets fixed to them in the vertical direction by stops. Each adjustable stop is made in the form of a threaded rod, the lower end of which is fixed to the support bracket (screwed into the support bracket), and the upper end through a spherical tip connected to the clamping plate in contact with its surface with the lower end surface of the corresponding wing panel SB. Moreover, the number of adjustable stops of each support node is equal to the number of panels in the wing of the SB, and the longitudinal axis of each stop is located in the plane of the contact panel of the wing of the SB. A rubber gasket is installed between the surface of the clamping bar of each adjustable stop and the lower end surface of the corresponding panel of the SB wing.

The proposed CO design has spacecraft attachment nodes that directly connect and hold the spacecraft PB on the LV support structure (CO locks), and nodes that provide spacecraft and spacecraft separation (CO pushers). In addition, the aircraft has attachment points located on the body of the aircraft (i.e., located outside the spacecraft satellites) under the wing of the satellites, on which the collapsed wing (wing panels with the help of hinged nodes are folded into a package - the transport inoperative position of the satin wings in the spacecraft launch site into the functioning orbit) and the SB wing fixed (with the help of fixation nodes) on the lateral surface of the SC SC.

The design of each fastener assembly (support node) of the SB, located on the body of the SB under the SB wing, consisting of adjustable stops interacting (in contact) with the lower end surfaces of the respective SB wing panels, provides wing fastening (support) in the longitudinal direction (in the direction of maximum axial overloads at the spacecraft launch site). Due to this, the SB wing is as if in a weightless state, the fixation nodes of the SB wing to the SC SC are unloaded and take only lateral loads acting on the SC, which are much smaller than the longitudinal ones. Due to the support under the SB wing, the loads on the CO locks are also reduced due to the eccentricity between the center of mass of the SB wing and the longitudinal axis of the SC SC (the SC wing is located on the side of the SC SC).

The support under the SB wing limits (decreases) its deflection and displacement with respect to the SC SC, which lead to the destruction (cracking) of photoelectric converters installed on the wing panels. The presence of fixation units for the SB wing on the SC satellite and support nodes under the SB wing located on the SB body significantly increase the rigidity of the SB wing installation on the SC compared to the technical solution, the prototype, in which the SB wing is fixed to the side of the SC satellite and on the CO adapter using fixation nodes (fasteners).

It should be noted that the location of each support assembly from below under the SB wing and the design of each support assembly (contact of the stops with each panel included in the wing) ensures the wing is supported under downward loads. And the loads directed upward, the support node does not perceive and therefore will not prevent the free separation of the spacecraft up from the efforts of the pushers interacting with the spacecraft satellite. Thus, unlike the technical solution - the prototype, the proposed attachment point of the CO does not have a mechanical connection that holds the wing (or panel) of the SB and, therefore, there is no influence of this support node located on the body of the SB under the wing of the SB on the functioning of the locks and pushers of the SB interacting with the spacecraft satellite.

To effectively increase the rigidity of the installation of the rolled-up SB wing on the PB KA and uniform loading of all its panels, contact of the stops with each panel is required, i.e. the number of stops in each support node should be equal to the number of panels in the wing of the SB. Otherwise, i.e. contact of the stops with not each wing panel of the SB will lead to uneven loading of both the panels themselves and the elements of their mutual fastening (hinge assemblies, interpanel connections, wing fixation nodes).

The location of the longitudinal axis of each stop in the plane of the SB wing panel in contact with it makes it possible to efficiently transfer the longitudinal inertial loads from each SB panel to the support unit and the CO body. If the longitudinal axis of each stop and the plane of the panel in contact with it (i.e., the axis of the stop leaves the plane of the panel) do not coincide, the stop is loaded with bending moment, and undesired loads from its plane would act on the SB panel at the points of contact with the stops.

The proposed design of threaded vertically adjustable stops, consisting of a threaded rod, the lower end of which is fixed to the support bracket, and the upper end through a spherical tip connected to the clamping bar, allows each stop to be brought into contact with the surface of its bar with the lower end of the corresponding wing panel of the SB on the accuracy of the manufacture and installation of individual SB panels in the wing. In real designs of multi-link wings, curvatures and non-flatness of the lower end surfaces (edges) of the panels are possible, which leads to their location at different heights when assembling into the wing. The design of the mount allows you to ensure the supply and guaranteed contact of the stops with the panels of the wing of the SB and to compensate for inaccuracies in the manufacture and installation of the wing of the SB on the PB SC. This is ensured by the rotation and tilt of the clamping plate on the spherical tip of the stop rod and the adjustment of the stops in height (in the vertical direction).

The installation between the surfaces of the strips and the lower end surfaces of the SB panels of rubber gaskets, crimped when adjusting the stops, allows to reduce shock and vibration shock loads acting on the CR when it is put into operation.

Thus, the proposed device separation system allows you to reduce the load on the fixation nodes of the wing of the SB to the SC CA and the locks CO fastening the SC to the LV, to reduce the deflection and movement of the wing of the SB relative to the SC SC, to increase the reliability of shock-free separation of the SC from the LV.

The figure 1 shows a General view of the spacecraft separation system with attachment points (support nodes) under the wing of the SB.

Figure 2 shows a view A according to figure 1.

The figure 3 shows a General view of the mount (see node B in figure 1).

The figure 4 shows a General view of the adjustable stop (see pos.11 in figure 3).

The figure 5 shows a section bb according to figure 4.

The system for separating the spacecraft from the LV support structure consists of the following devices (elements):

- spacecraft retention device on the LV support structure using locks mounted on the housing at the interface between the spacecraft PB and the LV support structure;

- a device for separating the spacecraft from the supporting structure of the launch vehicle - pushers that create pushing forces that act directly on the spacecraft's spacecraft when releasing the retaining connection in the locks;

- fasteners (support nodes) installed on the body of the CO under the SB wing, the panels of which are folded into a package.

The proposed design of the CO is installed between the supporting structure 1 of the launch vehicle and the spacecraft, including the PB 2 and the multi-link (multi-section) wing of the SB, fixed on the side, consisting of panels 3. The wing of the SB is in the transport (inoperative) position. Its panels 3 are folded into a package using hinge assemblies (not shown in the figures) and fixed (fixed) on the side surface of the PB 2 SC using fixation nodes 4. The CO consists of a housing 5, on which locks 6 and pushers 7 are installed. Locks 6 and pushers 7 WITH interact with PB 2 KA. The housing 5 is mounted on the supporting structure 1 of the launch vehicle and after separation of the spacecraft remains on the launch vehicle. WITH has support nodes 8 mounted on the body 5 under the panels 3 of the wing of the SB (see figures 1 and 2).

Each node 8 consists of a support bracket 9, mounted on the housing 5 WITH using threaded fasteners 10 (see figure 3). On the bracket 9 fixed stops adjustable in height (in the vertical direction) 11. The number of adjustable stops 11 on each support node 8 is equal to the number of panels 3 in the wing of the SB. Each stop 11 (see figure 4) consists of a threaded rod 12 and a clamping bar 13 fixed thereon. The clamping bar 13 is fastened to the threaded bar 12 through a spherical tip 14 made on the upper end of the threaded bar 12 using two plates 15 and 16 screws (see figure 5). In each plate 15, semicircular undercuts are made, providing coverage of the threaded rod 12 between the spherical tip 14 and the hexagon 17. The plates 15 are attached to the pressure plate 13 using screws 16 (two screws 16 on each plate 15, see figure 5, section B- C) and ensure that the spherical tip 14 does not fall out of the hole in the clamping bar 13. The size of the semicircular grooves in the plates 15 allows the clamping bar 13 to self-install relative to the lower end surface 21 of the corresponding panel 3 due to the possibility of rotation the clamp bar 13 on the spherical tip 14 of the threaded rod 12 of the stop 11. The lower threaded end of the rod 12 of the stop 11 is screwed into the support bracket 9, in the upper horizontal shelf of which the thread is made. The stop 11 is adjusted in height for the hexagon 17 made on the rod 12 by screwing (unscrewing) into the support bracket 9. When adjusting the stops 11, their lower threaded ends enter holes 18 made in the vertical wall of the support bracket 9 under each stop 11. After adjusting and installing the stops 11 in the desired position (until the contact of the clamping plate 13 of each stop 11 with the lower end surface 21 of the corresponding panel 3 and compression of the rubber strip 20 to the required size) is carried out and x locking relative to the bracket 9 with the help of nuts 19. On top of the surface of the clamping plate 13 of each stop 11 is fixed (for example, glued) a rubber gasket 20. The contact of the stops 11 with the panels 3 can be carried out directly with their lower end surfaces 21 or at the point of contact with the stops 11, corners 22 can be mounted on these surfaces. Thus, in the operating position of the CO device, the pressure bar 13 of each stop 11 is in contact with the lower end surface 21 (or corner 22) corresponding to the panel 3 of the SB wing through rubber gasket 20. The longitudinal axis of the stops 11 are located in the planes of the panels 3 of the SB wing in contact with them (see figure 1).

In the initial state of the separation system (in the spacecraft retention position on the LV), its elements are in the following position:

- the locks 6 are in a fixed (locked) position, interact with the lower part of the PB 2, holding the spacecraft on the supporting structure 1 PH;

- the pushers 7 are in the cocked state (for example, in the case of spring pushers, their springs are compressed by a certain amount necessary to create pushing forces on the spacecraft during its separation) and are also in contact with the lower part of the spacecraft PB 2;

- adjustable stops 11 of the supporting nodes 8 are in contact with the lower end surfaces 21 (or corners 22 mounted on these surfaces) of the panels 3 of the SB wing, providing direct support of the wing of the SB wing, the panels of which are folded into a bag and secured to the side of the PB 2 KA using fixation nodes four.

Thus, while holding the spacecraft on the LV, the mechanical connection between the spacecraft and the supporting structure 1 is carried out using locks 6 with the simultaneous support of the SB wing on the support node 8 mounted on the body 5 CO.

The device of the spacecraft separation system works as follows.

Before docking the spacecraft and the launch vehicle, all adjustable stops of the 11 attachment points of 8 СО are screwed into the support brackets 9. When the spacecraft is connected to the spacecraft, its PB 2 is mounted on the body of the 5 СО and fastened with the locks of 6 СО. Pushers 7 CO are adjusted and adjusted for specific triggering forces when the spacecraft is separated from the launch vehicle. At the same time, the wing panels 3 of the SB and the stops 11 do not contact each other, there is a gap between their lower end surfaces 21 (or corners 22) and the pressure bars 13 of the stops 11. Then, the stops 11 are unscrewed from the brackets 9 (for the hexagon 17 of the threaded rod 12) until the pressure strips 13 contact the corresponding panels 3 of the SB wing. The further unscrewing of the stops 11 compresses the rubber gaskets 20 to the required size (usually 50% of the thickness of the gasket) and locks the adjustable stops 11 in this position with the help of nuts 19.

At the spacecraft launch site, its construction is affected by linear inertial loads directed downward, which are transmitted through locks 6 (from PB 2) and support nodes 8 (from the SB wing) to the CO body 5 and then to the LV support structure 1. The SB wing, folded into a bag and secured to the side on the PB 2 with the help of fixation nodes 4 and supported from the support nodes 8 from below, has small deflections (deformations) and displacements. Possible alternating (vibrational) loads are compensated by installing rubber gaskets 20 and creating preliminary tightening forces of the adjustable stops 11. After unlocking (stripping) of the locks 6 and operation of the pushers 7, the spacecraft is separated from the supporting structure 1 PH. The design of the support nodes 8 WITH does not interfere with the separation of the spacecraft under the action of the efforts of the pushers 7, directed upward. When the spacecraft is separated from the LV, the panels 3 of the SB wing in the folded position freely move away from the stops 11 (between the pressure strips 13 of the stops 11 and the lower end surfaces 21 (or corners 22) of the panels 3, the contact disappears), mounted on brackets 9 mounted on the body 5 . The housing 5 with the support nodes 8 and the stops 11 mounted on them, remains on the supporting structure 1 of the pH. After separation of the spacecraft from the LV and after the stripping of the fixation nodes 4 of the panel 3 of the SB wing on the hinged nodes (not shown in the figures) from the folded (inoperative) position, they are deployed relative to the PB 2 to the working position.

The calculations performed by the authors, as well as bench tests as part of the developmental product, showed significant effectiveness of the proposed technical solution for mounting and separating the spacecraft from the launch vehicle.

The claimed device of the separation system compared with the prototype has significant differences and provides a reduction in the loads on the locks WITH, reducing deformations and movements of the wing of the SB during the launch of the spacecraft, increases the reliability of separation of the spacecraft from the launch vehicle.

Claims (2)

1. The separation system of the spacecraft, installed between the supporting structure of the launch vehicle and the spacecraft, containing a housing with locks and pushers placed on it, connecting the instrument unit of the spacecraft with the supporting structure of the launch vehicle and providing for their separation, equipped with mounting units located on the housing wing of the solar battery, consisting of panels mounted on the instrument unit of the spacecraft in the folded position, characterized in that each of the attachment points the wing of the solar battery is mounted on the housing of the separation system under the wing of the solar battery and consists of a support bracket with vertically adjustable stops fixed to it, made in the form of threaded rods, lower ends mounted on a support bracket, and the upper ends connected via spherical tips to pressure bars contacting their surfaces with the lower end surfaces of the respective panels of the wing of the solar battery, while the number of adjustable s of each of the attachment is equal to the number of panels in the wing of the solar battery, and the longitudinal axis of each adjustable stop is in the plane contacting with it the solar battery panel of the wing. 2. The spacecraft separation system according to claim 1, characterized in that a rubber gasket is installed between the surface of the pressure plate of each adjustable stop and the lower end surface of the corresponding panel of the wing of the solar battery.

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