CN108945519A - A kind of deployable satellite of integration based on hoop truss formula antenna - Google Patents
A kind of deployable satellite of integration based on hoop truss formula antenna Download PDFInfo
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- CN108945519A CN108945519A CN201810540310.0A CN201810540310A CN108945519A CN 108945519 A CN108945519 A CN 108945519A CN 201810540310 A CN201810540310 A CN 201810540310A CN 108945519 A CN108945519 A CN 108945519A
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- truss formula
- hoop truss
- formula antenna
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- satellite
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- 230000010354 integration Effects 0.000 title claims abstract description 16
- 230000007246 mechanism Effects 0.000 claims description 9
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- 238000013461 design Methods 0.000 abstract description 29
- 238000010586 diagram Methods 0.000 description 7
- 238000000034 method Methods 0.000 description 5
- 238000009434 installation Methods 0.000 description 4
- 238000005259 measurement Methods 0.000 description 3
- 108091092878 Microsatellite Proteins 0.000 description 2
- 230000005540 biological transmission Effects 0.000 description 2
- 238000004891 communication Methods 0.000 description 2
- 238000000354 decomposition reaction Methods 0.000 description 2
- 206010003591 Ataxia Diseases 0.000 description 1
- 206010010947 Coordination abnormal Diseases 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 230000002411 adverse Effects 0.000 description 1
- 238000013459 approach Methods 0.000 description 1
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- 238000010276 construction Methods 0.000 description 1
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/222—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles for deploying structures between a stowed and deployed state
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
Abstract
The deployable satellite of integration that the invention discloses a kind of based on hoop truss formula antenna, using payload platform integrated design, change the platform structure and load cabin structures in series mode of traditional " playing with building blocks " formula, around the structure of hoop truss formula antenna itself, the integrated design of platform load structure is realized using paralleling model, it can make full use of gathering and the expansion profiling characteristic of hoop truss formula antenna, design the hollow cylinder integration expandable type structure centered on hoop truss formula antenna, main body as platform structure, not only has good mechanical condition, maximally utilize the envelope of clear space, and it can realize modularized design, also has good interface with delivery；Meanwhile around with the distributed module of large-scale hoop truss formula antenna and control, system performance, such as mobility can be substantially improved, this provides a kind of new mentality of designing to by the design of the microwave satellite configuration of load of hoop truss formula antenna.
Description
Technical field
The invention belongs to technical field of satellite overall design, and in particular to a kind of integration based on hoop truss formula antenna
Deployable satellite.
Background technique
Satellite configuration refers to satellite integrally basic space frame and form, and satisfaction develops, emits, it is (right to run to return
Retrievable satellite) satellite whole life cycle in different desired overall spatial layout, shape and contour dimensions etc..Satellite structure
The design of shape is mainly by the constraint of several factors: payload, orbital characteristics, gesture stability mode and carrier rocket etc..Wherein,
Payload is its leading factor, and the configuration of satellite must be premised on the requirement for meeting payload.
In current satellite configuration design, it is still the situation of load Yu platform independent design, usually selects a certain
Microwave loaded antennas after the connection interface for specifying load and platform, is installed to platform knot based on this platform structure by platform
On structure.Such as the relevant patent of medium-and-large-sized satellite [a kind of configuration and its installation method CN102372092A of low rail remote sensing satellite],
[the satellite configuration CN105501471A for loading double-reflecting face Large deployable antenna] and microsatellite [octagon body packed battery
Battle array column micro satellite configuration CN102009746A] all embody this thought.The platform of satellite is deposited with microwave loaded antennas
At apparent interface, or even for the needs of the division of labor, platform and microwave loaded antennas are split.It is this to be defended for most of
Star is applicable, but for certain loaded antennas, structure type is changed, if still using traditional platform and load
The mode in cabin, so that the designing quality of satellite configuration is not high, if California bearing ratio is not high, structure efficiency is low, and optimization space is limited.Especially
It is to microwave load hoop truss formula antenna, and size is larger, is large-scale flexible body, and traditional approach is feed battle array on platform,
Antenna reflective face is put on display by large-scale unfolding mechanism.Antenna reflective face is very big compared to platform, and the inertia of whole star is also very
Greatly, the incoordination of large-scale load and small-sized platform, such as motor-driven, a series of classes can be brought by controlling chain-wales to adjust large-scale load
Equistability problem is directed toward like antenna beam.It controls difficulty greatly and control link is complicated (meeting load requirements by platform),
Especially large-scale expansion load.Have in the case where antenna and solar battery array both large scale compliant members at the same time
The high-precision attitude control for realizing whole star is the difficult point of satellite configuration design.Need in terms of overall configuration by integrated design come
Redundancy structure is reduced, Path of Force Transfer is shortened, reduces compliant member.Using loaded antennas, solar battery array as core, it is partially submerged into and defends
Star platform interior reinforces rigid support；The attitude measurements equipment such as star is quick, optical fibre gyro is directly installed on load main force support structure
On unified well with the benchmark and the benchmark of load that guarantee attitude measurement, shorten and turn between attitude measurement and load coordinate system
Error propagation chain is changed, the malformation and temperature control bring adverse effect of platform is avoided, improves the positioning accuracy of satellite.
For hoop truss formula antenna, due to the needs that SAR is scouted or is imaged, the bore of usual antenna is larger.Cause
This, examines payload platform integrated design, carries out Platform Structure Design around hoop truss formula antenna, whole star is made to accomplish that structure is tight
It gathers, lightly, improves the agility, mobility and stability of whole star, reduce the development difficulty of subsystem, save the cost.?
《Spacecraft antennas and beam steering methods for satellite communication
System " in (United States Patent, Patent Number:5642122), a similar idea is embodied, no
It crosses, emphasis is to be directed to communication need, rather than satellite configuration designs, and antenna system, antenna form, antenna structure and exhibition
It opens up in respect of greatly difference, it can be achieved that property is very poor.
Summary of the invention
In view of this, the deployable satellite of integration that the object of the present invention is to provide a kind of based on hoop truss formula antenna,
Have stable mechanical property when can guarantee satellite launch state, provide sufficient room for the layout and installation of satellite equipment, also
Carrier rocket clear space envelope can be made full use of.
A kind of deployable satellite of integration based on hoop truss formula antenna, including hoop truss formula antenna, and to receive
Surround centered on the hoop truss formula antenna held together under state and on the outside at least two equipment compartments of hollow cylindrical structure；
One of equipment compartment is as Cabin, for installing the feed battle array of the hoop truss formula antenna；The feed
Cabin is connect by unfolding mechanism with the hoop truss formula antenna；The reflecting surface of other equipment compartments and the hoop truss formula antenna
It is connected.
It further, further include Butt Section, for one end of the hollow cylindrical structure to be connected to carrier rocket
On interface.
Preferably, the quantity of the equipment compartment is 3-6.
It is conformal with the outer surface of equipment compartment and install preferably, the solar wing of the satellite selects flexible solar wing.
The invention has the following beneficial effects:
A kind of deployable satellite of integration based on hoop truss formula antenna of the invention, changes traditional " playing with building blocks " formula
Platform structure and load cabin structures in series fixed mode.Around the design feature of hoop truss formula antenna itself, using parallel connection
Mode realizes the deployable design of integration of platform load structure.It is embodied in centered on hoop truss formula antenna, periphery one is enclosed
Enclosed construction is formed, as load main structure and equipment compartment；Had stable mechanical property when both ensure that emission state, seals
Close the space to be formed and constitute " cabin " again, provide space for the layout of equipment and installation, and round section structure not only with week
Fringe truss formula antenna is conformal, also conformal with fairing of launch vehicle, can make full use of carrier rocket clear space envelope；In addition, circle
Shape section structure is also equipped with the extending space of short transverse, establishes good interface condition with carrier rocket.
Detailed description of the invention
Fig. 1 is the rounding state structural schematic diagram of integrated satellite of the invention；
Fig. 2 is the rounding state structural decomposition diagram of integrated satellite of the invention；
Fig. 3 is the in-orbit unfolded state schematic diagram of integrated satellite of the invention；
Fig. 4 is the design flow diagram of integrated satellite of the invention；
Wherein: 1- hoop truss formula antenna, 2- equipment compartment, the Butt Section 3-, 4- solar wing, 5- TT&C antenna, 6-GPS days
Line, 11- antenna reflective face, 12- antenna feed battle array, 13- unfolding mechanism.
Specific embodiment
The present invention will now be described in detail with reference to the accompanying drawings and examples.
For the microwave satellite with hoop truss formula antenna 1 for load, it is contemplated that it is collapsed and unfolded state is cylinder
Envelope, expansion process are expanded outward.Therefore, the present invention makes full use of load to collapse the cylindrical envelope feature with expansion, will go here and there
Join fixed platform structure and be changed to parallel deployable platform structure, realizes hoop truss formula antenna 1 and platform structure
The deployable Combined design of very high integrity.
According to device layout demand, design is improved to structure, increases interface, such as antenna feed battle array 12 and antenna-reflected
Face is in-orbit to need to keep relative positional relationship, and unfolding mechanism 13 is embedded into Cabin.Feed battle array different from the past is fixed on star
On ontology, antenna reflective face is unfolded by way of complicated；In-orbit expansion of the invention is incited somebody to action using antenna reflective face as main body
Feed battle array (being included in Cabin) is put on display.Solar battery array then uses flexible solar wing 4, flexible to be tightly attached to round cabin table
Face makes full use of star catalogue circular area.On the other hand, after in-orbit expansion, platform structure (cabin) and hoop truss formula antenna are anti-
11 close-coupled of face is penetrated, distributed uniform layout is conducive to the posture of entire antenna reflective face (big coasting body and big flexible body)
It is motor-driven.In addition, surrounding the distributed platform structure (cabin) of hoop truss formula antenna reflective face 11, platform is originally experienced hoop truss
11 space constraint of formula antenna reflective face is low, and star catalogue equipment (especially sensor and antenna) layout freedom is very high, changes
Influence of arrangement of traditional big load to star catalogue equipment short distance on chain-wales.
Attached drawing 1 is the satellite configuration schematic diagram of rounding state of the present invention, the circle formed centered on hoop truss formula antenna 1
Shape integration spliced platform structure in parallel, circular integrated platform structure are connect with calibre-changeable Butt Section 3, are established and delivery
Interface；
Attached drawing 2 is the satellite configuration decomposition diagram of rounding state of the present invention, and satellite is set by 1,3, hoop truss formula antenna
Standby 2,1, cabin Cabin and its unfolding mechanism 13, double solar wings 4, Butt Section 3 and several antennas composition；
Attached drawing 3 is that the configuration schematic diagram of unfolded state of the present invention forms one after the hoop truss formula antenna 1 of gathering is unfolded
It is a based on hoop truss formula structure, metal mesh as the large-scale antenna of reflecting surface, 3 equipment compartments 2 of antenna-reflected EDS maps, feedback
Source cabin then passes through unfolding mechanism 13 and antenna reflective face keeps certain positional relationship.
The deployable satellite configuration design of the integration that the present invention provides a kind of based on hoop truss formula antenna 1, can to
Hoop truss formula antenna 1 is that the satellite of main load carries out integration configuration design.Design cycle is configured as shown in figure 4, but this hair
It is bright detailed analysis not to be made to mission requirements, constraint condition (carrier rocket and track requirement etc.), microwave antenna system is directly selected
Select the hoop truss formula antenna reflective face system that the present invention considers.Mainly the features of the present invention is described in detail below.
Steps are as follows:
Overall configuration form.It is constrained according to fairing of launch vehicle, makes full use of the net enveloping space of radome fairing, counted
It calculates, specifies available inner diameter size.According to 11 collapsed diameter of hoop truss formula antenna reflective face, the envelope of hollow cylinder is divided.
In view of cylinder is a rock-steady structure, has preferable lateral mechanical characteristic, in emission state, entire integral structure
Lateral fundamental frequency can better adapt to the dynamics environment of carrier rocket.Therefore, hoop truss formula antenna reflective face 11 will be surrounded
Hollow cylinder platform structure as a whole carry out integrated design, guarantee the continuity in structure.Meanwhile considering week
Hollow cylinder Platform Structure Design can be expandable type, open circles by the configuration after the expansion of fringe truss formula antenna reflective face 11
Column platform structure can be divided into 3~6 modularization section structures, it is contemplated that modularization section structure, cabin device layout space, more
The convenience of the complexity, expansion that connect between section structure selects 3~4 cabins more appropriate, and the present invention uses 4 cabins
Body, as shown in Figure 1.
1 Method of Spreading Design of hoop truss formula antenna.Antenna feed battle array 12 and antenna reflective face operation on orbit need to keep opposite
Positional relationship, around 4 cabins of 11 reflecting surface of hoop truss formula antenna, 3 cabins and hoop truss formula antenna reflective face 11
It is unfolded together, 1 cabin is set as antenna feed cabin, feed battle array is embedded into Cabin, in-orbit expansion integrally opens up Cabin
Out.According to the required precision of hoop truss formula antenna 1, hinge unfolding mechanism 13 is designed.Hinge unfolding mechanism 13 needs while expiring
Foot transmitting when locking, enter the orbit after be unfolded.
With the Interface design of delivery.With the hollow cylinder of hoop truss formula antenna reflective face 11 structure as a whole,
Using round calibre-changeable Butt Section 3, as the transition connecting structure between carrier rocket and satellite, realize that carrier rocket circle connects
Mouth is docked with bottom cylindrical face interface.The hollow cylinder lower end surface of monoblock type is connected with 3 upper surface of Butt Section, finally concentrates on phase
On 8 points handed over (each cabin two sides and the internal whole partition for having a longitudinal direction), it is transmitted to Butt Section 3.Therefore, hollow cylinder
The connection of lower end surface and 3 upper surface of Butt Section is crucial, is a kind of stable power transmission form using 8 power transmissions.It can not only expire
Sufficient emission state, and be able to satisfy enter the orbit after separate.
Solar wing layout designs.Solar wing uses flexible structure, and emission state utilizes outside hollow cylinder, as
The installation stationary plane of solar wing 4.Tie point can directly utilize the vertical and horizontal frame structure of 2 body of equipment compartment, rigidity herein
Higher, mechanical environment is preferable, is conducive to emission state.Flexible solar wing 4 is first unfolded, locks after entering the orbit, flexible solar wing 4 and cabin
Connection relationship is formed between body.
Device layout.Equipment mainly depends on platform structure cabin, 4 cabins that hollow cylinder platform structure divides in star
Body, each cabin can carry out modularized design, wherein 1 as Cabin (the modularization cabin of antenna feed battle array 12, be laid out feed
Battle array and associated electronic device), power module cabin, control module cabin and propulsion die cabin etc. or 1 can be arranged in remaining according to demand
Modularization inside cabin.Star catalogue equipment is then since platform structure cabin (gathering and unfolded state) is anti-by hoop truss formula antenna
It penetrates that 11 space constraint of face is low, and layout freedom is very high, especially sensor and TT&C antenna 5, GPS antenna 6 etc., can flexibly use
The device layout that antenna etc. requires visual field is passed in sensor and observing and controlling number.
The hollow cylindrical body structure established centered on hoop truss formula antenna 1 is optimized.By aforementioned
Step primarily determines the configuration layout of satellite, then establishes the model via dynamical response of whole star, carries out the Structural Static of emission state
Mechanics and dynamic analysis obtain whether rigidity of integral structure etc. meets the requirements especially by dynamic analysis, main biography
Whether power path design is reasonable, then optimizes.
In conclusion the above is merely preferred embodiments of the present invention, being not intended to limit the scope of the present invention.
All within the spirits and principles of the present invention, any modification, equivalent replacement, improvement and so on should be included in of the invention
Within protection scope.
Claims (4)
1. a kind of deployable satellite of integration based on hoop truss formula antenna, which is characterized in that including hoop truss formula antenna,
And surround centered on the hoop truss formula antenna under rounding state and on the outside at least two of hollow cylindrical structure
Equipment compartment；
One of equipment compartment is as Cabin, for installing the feed battle array of the hoop truss formula antenna；The Cabin is logical
Unfolding mechanism is crossed to connect with the hoop truss formula antenna；The reflecting surface of other equipment compartments and the hoop truss formula antenna is solid
Even.
2. a kind of deployable satellite of integration based on hoop truss formula antenna as described in claim 1, which is characterized in that also
Including Butt Section, for one end of the hollow cylindrical structure to be connected to the interface of carrier rocket.
3. a kind of deployable satellite of integration based on hoop truss formula antenna as described in claim 1, which is characterized in that institute
The quantity for stating equipment compartment is 3-6.
4. a kind of deployable satellite of integration based on hoop truss formula antenna as described in claim 1, which is characterized in that institute
The solar wing for stating satellite selects flexible solar wing, conformal with the outer surface of equipment compartment and install.
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| CN201810540310.0A CN108945519A (en) | 2018-05-30 | 2018-05-30 | A kind of deployable satellite of integration based on hoop truss formula antenna |
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| CN201810540310.0A CN108945519A (en) | 2018-05-30 | 2018-05-30 | A kind of deployable satellite of integration based on hoop truss formula antenna |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112257191A (en) * | 2020-12-23 | 2021-01-22 | 中国人民解放军国防科技大学 | Load platform integrated microsatellite thermal control subsystem optimization method and system |
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| JPH05213283A (en) * | 1992-02-04 | 1993-08-24 | Nec Corp | Structure of artificial satellite |
| US20090057492A1 (en) * | 2007-08-28 | 2009-03-05 | Harris Mark A | Space vehicle having a payload-centric configuration |
| CN101834350A (en) * | 2009-03-13 | 2010-09-15 | 中国科学院国家天文台 | Passive cable-leading-in-cavity connection mechanism of large-aperture radio telescope |
| CN102009746A (en) * | 2010-11-08 | 2011-04-13 | 航天东方红卫星有限公司 | Octagonal battery-equipped array upright post micro satellite configuration |
| CN104908979A (en) * | 2015-05-11 | 2015-09-16 | 上海宇航系统工程研究所 | Flexible solar wing compaction and release device |
| CN105035358A (en) * | 2015-07-31 | 2015-11-11 | 上海卫星工程研究所 | In-orbit expansion-type satellite structure |
| CN104443431B (en) * | 2014-10-23 | 2017-08-29 | 上海卫星工程研究所 | Triangle satellite configuration |
-
2018
- 2018-05-30 CN CN201810540310.0A patent/CN108945519A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH05213283A (en) * | 1992-02-04 | 1993-08-24 | Nec Corp | Structure of artificial satellite |
| US20090057492A1 (en) * | 2007-08-28 | 2009-03-05 | Harris Mark A | Space vehicle having a payload-centric configuration |
| CN101834350A (en) * | 2009-03-13 | 2010-09-15 | 中国科学院国家天文台 | Passive cable-leading-in-cavity connection mechanism of large-aperture radio telescope |
| CN102009746A (en) * | 2010-11-08 | 2011-04-13 | 航天东方红卫星有限公司 | Octagonal battery-equipped array upright post micro satellite configuration |
| CN104443431B (en) * | 2014-10-23 | 2017-08-29 | 上海卫星工程研究所 | Triangle satellite configuration |
| CN104908979A (en) * | 2015-05-11 | 2015-09-16 | 上海宇航系统工程研究所 | Flexible solar wing compaction and release device |
| CN105035358A (en) * | 2015-07-31 | 2015-11-11 | 上海卫星工程研究所 | In-orbit expansion-type satellite structure |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN112257191A (en) * | 2020-12-23 | 2021-01-22 | 中国人民解放军国防科技大学 | Load platform integrated microsatellite thermal control subsystem optimization method and system |
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