RU2374148C2 - Spacecraft modular design - Google Patents

Abstract

FIELD: space engineering.

SUBSTANCE: invention relates to spacecraft designs, particularly to artificial satellites. Modular structure comprises bearing structure representing horizontal package consisting of vertical packs of unified crates (1, 1a, 1b, 2, 3). Crates (1, 1a, 1b) of one type are intended for VME or PCI form-factor. Crates of the other type (2, 3) serve to accommodate optional instruments and units, for example storage batteries. Crate height can be subject to quantity of identical-thickness electronic modules located along crate height, and to module form-factor spacing in height. Crate wall thickness ensures required thickness of spacecraft in transfer of loads from carrier, as well as radiation protection of electronic modules. Crates are jointed together with the help of brackets that support elements (12) incorporated with the system intended for separation from carrier. At least one crate of vertical package (1a, 1b) can be arranged via thermal insulating gaskets with their total thickness being multiple to aforesaid form-factor spacing in height.

EFFECT: reduced weight and costs, modular design.

3 cl, 7 dwg

Description

The invention relates to space technology and, in particular, to the design of modular spacecraft (satellites).

From the patent literature there are many designs of modular spacecraft, consisting, as a rule, of the supporting structure and various devices and assemblies installed on it. In this case, both the supporting structure (body) and the devices can be modular.

For example, in patents [1, 2, 3] and in applications [4, 5], modular load-bearing structures are presented, and in patents [6, 7, 8] unified racks with instrument modules are presented.

It is also known [9] that there is a tendency to use commercially available electronic devices and components in space technology, COTS (Commercial Of The Shelf). Such components are ten times cheaper and 2-3 generations more perfect than components traditionally used in space technology. The main problems of using such components in space technology are vibration shock resistance, heat dissipation and radiation resistance.

Problems with vibration shock resistance and heat dissipation are easily solved in IEEE 1101.10 and IEEE 1101.2 (Eurocard) standards, especially for the 3U form factor, and subracks made in these standards are widely used in military equipment. One of the advantages of using the mentioned standards is the use of modern bus architectures of the VME, CompactPCI, CompactPCI Express, etc. standards, which significantly reduces the number of interconnect cables and allows the use of commercially available cards in these standards with minimal modifications.

The problem with radiation resistance of commercially available components, in particular processors, main memory and mass memory, is solved both at the hardware-software level and at the constructive level. At the software and hardware level, problems with single / multiple errors from the passage of fast protons and heavy particles are solved. The problem with the accumulated radiation can be solved only by compaction of the layout and shielding of electronic components by structural elements.

For example, devices made in the DORA architecture [10] have a dense internal layout with an aluminum case wall thickness of 10-15 mm.

In all known solutions, the supporting structure of the apparatus also participates in radiation protection, but the integral thickness of the screen is not more than 2-4 mm aluminum, especially when using three-layer honeycomb panels.

An increase in the density of the device’s layout in inventions [1, 2, 3, 4, 5] and the placement of radiation-critical devices behind the screen from non-critical, for example, gyrodines or rechargeable batteries, can bring the integral screen thickness for the most critical devices to the required 10-15 mm . But at the same time, the modular architecture of the device, as a rule, is no longer necessary, since modularity inevitably implies certain restrictions on freedom of layout and some redundancy, including the volume of the device. The situation can be corrected by introducing additional screens and increasing the thickness of the walls of the devices, but this will lead to an increase in mass.

The situation in the invention [6], adopted by the authors as a prototype, is somewhat better. In this invention, there is a supporting structure with fastening elements of devices and units of service and target equipment, providing rigidity of the device and transferring loads from the carrier, and unified instrument racks (bays) placed on it, in the slots of which are placed unified electronic modules (cards) of service and target equipment. The placement of unified electronic modules in the crates significantly increases the density of the electronic equipment layout, which allows relatively low costs to organize radiation protection by increasing the thickness of the walls of the crates. The disadvantage of this design is that the excess wall thickness of the subracks is used only as a radiation screen, and excess weight leads to an increase in the mass of the supporting structure and the apparatus as a whole. Another disadvantage is that the height of the subracks and their capacity are rigidly determined by the height of the supporting structure. Accordingly, either the supporting structure of the apparatus and the racks will have to be developed again for each specific mission, with a specific composition of the equipment, or the supporting structure of the apparatus and the racks for some missions will be redundant and "underloaded" with electronic equipment.

The objective of the invention is to create a modular design of the spacecraft, which simultaneously solved the problems of both the modularity of the supporting structure of the spacecraft and the modularity of electronic equipment for office and target equipment, while reducing the mass of the device and its cost.

This problem is solved by the fact that in the modular design of the spacecraft containing the supporting structure, which provides the spacecraft rigidity when transferring loads from the carrier, and standardized crates with electronic modules of service and target equipment, having walls with a thickness necessary for radiation protection of electronic modules, the supporting structure of the spacecraft made as a horizontal package of vertical racks.

At least two vertical packages can have the same height, and the height of the racks obey the formula:

H = (N + 1)

where H is the height of the rack;

N is the number of electronic modules of unit thickness placed vertically in the crate;

h is the vertical step of the electronic module form factor.

At least one rack in a vertical package can be installed through heat-insulating gaskets, the total thickness of which is a multiple of the indicated step h.

Further, the invention is disclosed in more detail using graphic materials, in which figure 1 shows the layout options of spacecraft; Figure 2 - supporting structure of one of the options; Figure 3 is a variant of a uniform row of racks; Figure 4 - design of the rack with standard electronic modules of the same form factor; Figure 5 - design of the rack with electronic modules of different form factors; 6 is a design of the rack for non-standard elements and devices; 7 - three layout options for a vertical package of racks with the same total height. For simplicity of perception, all images do not show non-essential elements, for example, fasteners, adapter brackets, connectors, etc.

Figure 1 shows two options for the modular design of the spacecraft: a high-resolution remote sensing satellite (Earth), and an ultra-high-resolution satellite.

The simplest version of the microsatellite (not shown) is also possible, for example, a vertical package of 1-2 crates, with the installation of all electronic modules of service and target equipment, including flat battery blocks made in the form factor of the electronic module, in the slots of the crates.

It should be noted that in this invention modularity leads to the fact that there is no technical need to strictly separate the design of the device into traditional modules (compartments) of service and target equipment (exceptions see below). Almost all electronics (for example, in the CompactPCI Express standard), regardless of accessory, are located in unified racks 1. Some elements and devices that cannot be reduced to the CompactPCI form factor and which do not have special layout restrictions (unlike, for example, gyrodynes), are placed in crates 2 (battery pack) and 3 (capacitor bank), which have the same external and connecting dimensions as crates 1.

Outside on the racks are placed optical modules (telescopes) 4, 5 with star sensors 6 mounted directly on the telescopes, gyrodynes 7, electromagnetic actuators 8 and ablative pulsed plasma engines 9. Other equipment, for example, temperature control radiators, solar panels, is also installed on the racks et al. (not shown).

The actual supporting structure of the apparatus is shown in FIG. 2, where the vertical packages of the racks 1, 2, 3 and the vertical package of the racks 1a and 1b are shown in an exploded view. Between themselves, the vertical packages are connected in a horizontal package with the help of brackets 10 and 11. On the brackets 10 there are installed elements 12 of locks for fastening to the carrier. The bracket 11 has landing sites 13 for mounting the telescope 5 (Figure 1).

Figure 3 shows a variant of a unified series of standard sizes of crates. At the same time, a number of standard sizes are shown on lines A and B, and on line B vertical packages of two racks are shown, corresponding to the height of the racks on line B.

Figure 4 shows the design of the crate form factor 3U (100 × 160 mm) CompactPCI Express. The rack consists of side walls 14 and 15, the front and rear walls 16 and 17, the upper and lower walls 18 and 19, and the cross-board 20. Electronic modules (cards) 21 and 22 are installed in the slots of the rack. The rear wall has openings 23 for installation Power, I / O, CompactPCI Express, and other connectors (not shown). The front wall 16 may have holes for connectors that are impossible or impractical to output through the cross-board 20 and the rear wall 17, for example, high-frequency connectors (not shown). Walls 14, 15, 18, 19 have fastening elements 24 (threaded holes) for installing non-standard devices and components from the outside of the crate. The side walls 14 and 15 have structural elements (landing pads) 25 for transmitting external loads. Figure 5 shows a design variant of a universal rack supporting both standard modules 21 and 22 of the 3U form factor and modules 26 of a non-standard form factor 210 × 160 mm. Cross-board 27 is designed to support various combinations of modules of both form factors with the corresponding intermediate guide bracket 28. Figure 6 shows the rack 2 (Figure 2) for non-standard devices and assemblies, in this case the battery pack 29 .

It was noted above that in this architecture of the device there is no technical need to divide the design into traditional modules (compartments) of service and target equipment. However, for organizational reasons, it may be necessary to separate the modules of any system of the device into a separate crate. In addition, it may be necessary to provide a special thermal regime for any electronic modules.

7 (A) shows a rack 1 of height H in increments of the form factor h. It contains N electronic modules 23. The wall thickness S for radiation protection should be at least 5-6 mm, which, as a rule, exceeds the thickness required for strength reasons for such a supporting structure of the apparatus. Of course, the wall thickness S, together with other structural elements, should provide the necessary threshold for integrated radiation protection for most electronic components, which does not exclude the individual protection of specific components such as processors and, especially, operational and non-volatile mass memory. Such protection can be performed directly on electronic modules in the area where specific components are located using, for example, tungsten plates.

Obviously, from layout considerations, the most optimal is a supporting structure in which at least two, or better, all vertical rack packages have the same height. 7 (B) shows a package of crates 1a and 1b with a total height equal to the sum of H1 and H2 and equal to the height H of crate 1.

In this case, the height (H) of each rack obeys the formula:

H = (N + 1)

where N is the number of electronic modules of unit thickness placed in the crate in height, h is the pitch of the form factor of the electronic module in height.

In general, the height (H _Σ ) of the rack package obeys the formula:

H _Σ = (N ₁ + K ₁ ) · h + ... + (N _i + K _i ) · h,

where K _i = 0,1,2 ... is an integer coefficient depending on the design of the rack (in this case, for all racks, K = 1). At K = 0, the crate has no upper and lower covers and can be placed in the middle of the package (not shown).

Thus, it is quite flexible (within the selected range of sizes) to select combinations of racks to form the optimal load-bearing structure of the apparatus.

7 (B) shows a vertical package of crates 1c and 1g with a total height also equal to the height H of crate 1. At that, crate 1c is thermally insulated with gaskets 30 made of a material with low thermal conductivity and with a total height h1 and h2 equal to h . Thus, in conjunction with the installation of individual radiators, heaters and EVTI (not shown), it is possible to provide a special thermal regime for a rack 1v.

The use of a unified series of subracks providing the necessary radiation protection to electronic equipment based on COTS components and performing the function of the supporting structure of the spacecraft will allow achieving the following advantages.

Modularity and unification will significantly accelerate the development, manufacture and testing of the apparatus, which, combined with the use of COTS components, will reduce its cost, and the absence of a traditional supporting structure will reduce the mass of the apparatus.

Thus, all the tasks of the present invention are completed.

Literature

1. Patent US 6053454.

2. Patent US 6206327.

3. Patent US 6220548.

4. Application US 2006/0016935.

5. Application US 2007/0029446.

6. Patent US 5518209 (prototype).

7. Patent US 6298289.

8. Patent US 6721658.

9. Article "An Integrated Approach with COTS Creates Rad-Tolerant Single Board Computer for Space", COTS Journal, December 2003, http://www.cotsjournalonline.com/.

10. Vlasov S.F., Vlasov F.S., Chefranov M.I. Technology for the development and production of durable fail-safe and radiation-resistant equipment. III International Conf. - exhibition “Small satellites. New technologies, miniaturization. Areas of effective use in the XXI century ”, May 27-31, 2002: Sat. doc. - M. - 2002. - Prince III.

Claims (3)

1. The modular design of the spacecraft, containing the supporting structure, providing the rigidity of the device when transferring loads from the carrier, and standardized crates with electronic modules of service and target equipment, having walls with a thickness required for radiation protection of electronic modules, characterized in that the supporting structure is made like a horizontal bag made up of vertical racks. 2. The modular design according to claim 1, characterized in that at least two vertical packages of racks have the same height, and the heights (H) of the racks obey the formula:
H = (N + 1)
where N is the number of electronic modules of unit thickness placed in the crate in height, h is the pitch of the form factor of the electronic module in height. 3. The modular design according to claim 2, characterized in that at least one rack in a vertical package is installed through heat-insulating gaskets, the total thickness of which is a multiple of the indicated step h.

Images