WO2024087465A1 - 一种星箭载一体化飞行器 - Google Patents

一种星箭载一体化飞行器 Download PDF

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Publication number
WO2024087465A1
WO2024087465A1 PCT/CN2023/081471 CN2023081471W WO2024087465A1 WO 2024087465 A1 WO2024087465 A1 WO 2024087465A1 CN 2023081471 W CN2023081471 W CN 2023081471W WO 2024087465 A1 WO2024087465 A1 WO 2024087465A1
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Prior art keywords
rocket
satellite
carrier
stage
integrated
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PCT/CN2023/081471
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English (en)
French (fr)
Inventor
曹喜滨
郭金生
邱实
吴凡
岳程斐
魏承
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哈尔滨工业大学
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Publication of WO2024087465A1 publication Critical patent/WO2024087465A1/zh

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/50—On board measures aiming to increase energy efficiency

Definitions

  • the present disclosure relates to the technical field of aerospace structure design, and more specifically to a satellite-rocket integrated aircraft.
  • solid carriers such as solid carrier rockets
  • the solid carriers have limited capacity in terms of the weight of the payload of the satellites they carry.
  • the capacity of the solid carrier in a sun-synchronous orbit at a distance of 500km is about 200 ⁇ 1200Kg
  • the weight of the payload in orbit accounts for about 10 ⁇ 30%, which seriously restricts the space response capability to sudden events. Therefore, in order to improve the space rapid response capability of solid carriers, it is of great strategic significance to develop an integrated system of satellites and solid carrier rockets (hereinafter referred to as "satellites and rockets").
  • the integrated system has the characteristics of rapid development, rapid integration, rapid testing, rapid launch and rapid orbit entry, and can be applied to various sudden emergency situations (for example, sudden natural disasters and failures of communication systems, etc.) to realize the rapid launch and space deployment of satellites, obtain sudden event information in a timely manner, minimize losses and organize to fight against sudden events.
  • sudden emergency situations for example, sudden natural disasters and failures of communication systems, etc.
  • the integrated rocket-satellite vehicle requires a lightweight structure. This can be achieved by reducing the weight of the payload carried by the launch vehicle.
  • the rocket and satellite in the integrated vehicle share some duplicate equipment and structures, which greatly reduces the redundant weight of the vehicle entering orbit.
  • the payload of the satellite cabin in the existing integrated rocket-satellite vehicle after entering orbit is relatively low, which affects the further reduction of the vehicle's takeoff weight.
  • relevant technical research has been carried out on the integrated rocket-satellite vehicle system, but there are still some problems, specifically:
  • the payload, propulsion system, and electronic system of the aircraft have a low degree of modularity and do not utilize mass production and testing.
  • the embodiments of the present disclosure are intended to provide an integrated satellite-rocket aircraft that can enhance the carrying capacity of the carrier, increase the payload weight ratio of the satellite entering orbit, and at the same time, improve the modularity of the aircraft system to achieve mass production and testing.
  • the embodiment of the present disclosure provides a satellite-rocket integrated aircraft, the satellite-rocket integrated aircraft includes a satellite and a carrier, the satellite includes a payload, an instrument cabin, a power control system and a solar panel installed on the outside of a tank in the power control system, the carrier includes a carrier sub-stage first-stage rocket, a carrier sub-stage second-stage rocket, a carrier sub-stage third-stage rocket and a carrier sub-stage final-stage rocket; wherein,
  • the satellite and the carrier sub-stage final rocket of the carrier share the tank in the power control system to form a satellite-rocket integrated aircraft;
  • the satellite-rocket integrated aircraft is installed inverted on the transition section between the carrier-stage three-stage rocket and the carrier-stage final-stage rocket to be connected to the carrier through the payload.
  • the embodiments of the present disclosure provide a satellite-rocket integrated aircraft; the satellite and the carrier sub-stage final rocket of the carrier share a propulsion system to form a satellite-rocket integrated aircraft, and then the satellite-rocket integrated aircraft is installed on the transition section to achieve an integrated design of the satellite, carrier and payload, and at the same time, by controlling the weight ratio of the payload to increase the weight ratio of the payload when the satellite enters orbit, so that the satellite-rocket integrated aircraft can better realize the observation of sudden events.
  • FIG1 is a schematic structural diagram of a satellite-rocket integrated aircraft provided by an embodiment of the present disclosure
  • FIG2 is a partial enlarged view of portion A in FIG1 ;
  • FIG3 is an exploded schematic diagram of a satellite structure provided by an embodiment of the present disclosure.
  • FIG4 is a schematic diagram of the connection between the satellite and the carrier sub-stage final rocket in a conventional technical solution
  • FIG5 is a schematic diagram of the connection between the main structure of the payload and the mounting plate of the instrument cabin provided by an embodiment of the present disclosure
  • FIG6 is a schematic cross-sectional view along line B-B in FIG5 ;
  • FIG7 is a schematic diagram of the structural assembly of a payload and an instrument cabin provided by an embodiment of the present disclosure
  • FIG8 is a schematic diagram of heat transfer between a payload and an instrument cabin provided by an embodiment of the present disclosure
  • FIG9 is a schematic diagram of the structural assembly of a satellite provided by an embodiment of the present disclosure.
  • FIG. 10 is a schematic diagram of the structure of a power control system provided in an embodiment of the present disclosure.
  • the satellite-rocket integrated aircraft 1 includes a satellite 11 and a carrier 12.
  • the satellite 11 includes a payload 111, an instrument cabin 112, a power control system 113, and a solar panel 114 installed on the power control system 113.
  • the carrier 12 includes a carrier-stage first-stage rocket 121, a carrier-stage second-stage rocket 122, a carrier-stage third-stage rocket 123, and a carrier-stage final-stage rocket 124.
  • the satellite 11 and the carrier sub-stage final rocket 124 of the carrier 12 share the tank 1131 in the power control system 113 to form a satellite-rocket integrated aircraft 13;
  • the satellite-rocket integrated vehicle 13 is installed inverted on the transition section 1231 between the carrier-stage third-stage rocket 123 and the carrier-stage final-stage rocket 124 to be connected to the carrier 12 through the payload 111.
  • the satellite 11 and the carrier sub-stage final rocket 124 are respectively separated from the docking ring 43 and the satellite support cabin 44 on the carrier sub-stage final rocket 124 under the action of the satellite thruster 41 and the final stage thruster 42, thereby separating the satellite 11 from the carrier sub-stage final rocket 124 of the carrier 12 to form a split launch mode for the carrier 12 and the satellite 11.
  • the carrier sub-stage final rocket 124 of the carrier 12 and the satellite 11 share a propulsion system to form an integrated launch mode for the carrier 12 and the satellite 11.
  • a main columnar structure 1111 is provided in the payload 111, and a plurality of satellite-rocket connection points 1112 are provided on the outer periphery of the upper end of the main structure 1111 for connecting the satellite-rocket integrated aircraft 13 with the transition section 1231.
  • the connections between the multiple satellite-rocket connection points 1112 and the transition section 1231 are respectively provided with vibration isolators 52 shown in FIG. 5 to reduce the influence of the active section vibration of the carrier 12 on the payload 111 .
  • the main structure 1111 and the mounting plate 31 of the instrument cabin 112 are connected by bolts, and a vibration isolator 52 is provided at the connection between the main structure 1111 and the mounting plate 31 of the instrument cabin 12 to reduce the influence of the vibration of the instrument cabin 112 on the payload 111.
  • a limiting bushing 61 is further provided between the bolt 51 and the vibration isolator 52 .
  • the instrument bay 112 is used to install aircraft instruments 32 , including electronic equipment, power controllers, batteries and other electronic equipment, and the satellite support bay 44 .
  • a multi-layer heat-conducting film 81 is provided between the main structure 1111 and the mounting plate 31 of the instrument cabin 112. It can be understood that in the specific implementation process, the number of layers of the heat-conducting film 81 can be adjusted according to different thermal control requirements to achieve different heat transfer or heat insulation effects.
  • the main structure 1111 and the tank 1131 in the power control system 113 are connected by bolts, and a vibration isolator 52 is provided at the connection between the main structure 1111 and the tank 1131 in the power control system 113 to reduce the influence of the vibration of the power control system 113 on the payload 111.
  • a mounting hole 311 is provided on the mounting plate 31, so that the base 11311 of the tank 1131 can be connected to the main structure 1111 of the payload 111 by bolts through the above-mentioned mounting hole 311.
  • a bolt hole 11312 is provided on the base 11311 of the tank 1131.
  • Figure 9 shows a schematic diagram of the structure of the satellite 11 after assembly.
  • the carrier 12 directly transmits force to the satellite 11 through the satellite-rocket connection point 1112 on the payload 111, achieving the shortest overall force transmission path of the satellite-rocket integrated aircraft 1.
  • the carrier 12 and the satellite 11 are integrated into one design, which realizes the lightweight design of the satellite-rocket integrated aircraft 1 and increases the weight proportion of the payload 111 of the orbiting satellite 11. Specifically, the weight proportion of the remaining systems in the satellite 11 is 30%, and the weight proportion of the payload 111 can reach 70%.
  • the main structure 1111 of the payload 111 is the main load-bearing structure, and therefore a reinforced design is performed during the specific implementation process to better transmit force.
  • the power control system 113, the instrument cabin 112 and the payload 111 can be independently assembled and tested. It can be understood that the payload 111, the instrument cabin 112 and the power control system 113 in the satellite 11 are modularly designed to facilitate independent assembly and testing. Specifically, the modular design of the payload 111, the instrument cabin 112 and the power control system 113 realizes rapid mass production and independent mass testing, which can shorten the assembly and testing time by 70%.
  • a vibration isolator 52 is provided between the tank 1131 of the power control system 112 and the mounting plate 31 of the instrument cabin 112 to reduce the effect of the liquid sloshing in the tank 1131 on the payload 111.
  • a flywheel 101, a magnetic torquer 102, a track control engine 103, a plurality of attitude control engines 104 and a plurality of high-pressure gas cylinders 105 are also installed on the outside of the tank 1131 of the power control system 113; wherein the high-pressure gas cylinders 105 are symmetrically arranged on the bottom surface of the tank 1131; the track control engine 103 is arranged between the high-pressure gas cylinders 105 and ejects gas outward to generate propulsion; the attitude control engine 104 is symmetrically arranged on the outside of the high-pressure gas cylinders 105 and ejects gas outward to generate propulsion.
  • a vibration isolator 52 is provided between the tank 1131 and the flywheel 101 to reduce the influence of the micro-vibration of the flywheel 101 on the payload 111 .
  • the vibration isolator 52 can be a T-shaped rubber vibration isolator, or other types of rubber vibration isolators can be used according to actual needs; of course, the vibration isolator 52 can be of different sizes according to different application scenarios.
  • the instrument cabin 112 is installed on the payload 111 structure to achieve satellite-borne integration, and the satellite 11 and the carrier sub-stage final rocket 124 of the carrier 12 share a propulsion system to form a satellite-rocket integrated aircraft 13, and then the satellite-rocket integrated aircraft 13 is installed on the transition section 1231 between the carrier sub-stage third-stage rocket 123 and the carrier sub-stage final rocket 124 to achieve an integrated design of the satellite 11, the carrier 12 and the payload 111, and at the same time, by controlling the weight proportion of the payload 111, the weight proportion of the payload 111 when the satellite 11 enters orbit is increased, so that the satellite-rocket integrated aircraft 1 can better realize the observation of sudden events.
  • the embodiments of the present disclosure provide a satellite-rocket integrated aircraft; the satellite and the carrier sub-stage final rocket of the carrier share a propulsion system to form a satellite-rocket integrated aircraft, and then the satellite-rocket integrated aircraft is installed on the transition section to achieve an integrated design of the satellite, carrier and payload, and at the same time, by controlling the weight ratio of the payload to increase the weight ratio of the payload when the satellite enters orbit, so that the satellite-rocket integrated aircraft can better realize the observation of sudden events.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Remote Sensing (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
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Abstract

本公开实施例公开了一种星箭载一体化飞行器，属于航天结构设计技术领域；所述星箭载一体化飞行器包括卫星和运载器，所述卫星包括有效载荷、仪器舱、动力控制系统及安装于所述动力控制系统中的贮箱外部上的太阳能帆板，所述运载器包括运载子级一级火箭，运载子级二级火箭，运载子级三级火箭和运载子级末级火箭；其中，所述卫星与所述运载器的运载子级末级火箭共用所述动力控制系统中的贮箱以形成星箭一体化飞行器；所述星箭一体化飞行器被倒置地安装于所述运载子级三级火箭与所述运载子级末级火箭之间的过渡段上，以通过所述有效载荷与所述运载器进行连接。本发明实施例提供的星箭载一体化飞行器能够提升卫星的运载能力和有效载荷的重量占比。

Description

一种星箭载一体化飞行器 相关申请的交叉引用
本申请要求于2022年10月24日提交中国专利局、申请号为2022112996662，发明名称为“一种星箭载一体化飞行器”的中国专利申请的优先权，该中国专利申请的内容在此引入本申请作为参考。
技术领域
本公开涉及航天结构设计技术领域，更具体地涉及一种星箭载一体化飞行器。
 背景技术
为实现对突发性事件的观测，卫星多搭载固体运载器发射，尽管固体运载器，例如固体运载火箭，能够对突发性事件做出快速响应，但是固体运载器在所运载卫星的入轨有效载荷重量方面的能力有限，具体来说，在距离500km的太阳同步轨道上固体运载器的能力大约为200~1200Kg，入轨有效载荷的重量占比约为10~30%，严重制约了应对突发性事件的空间响应能力。因此，为了提升固体运载器空间快速响应能力，发展卫星与固体运载火箭（以下简称“星箭”）一体化系统具有重要的战略意义。一体化系统具有快速研制、快速集成、快速测试、快速发射和快速入轨等特点，可以应用于各种突发应急状况（例如，自然灾害突发和通信系统发生故障等），以实现卫星的快速发射和空间部署，及时获取突发性事件信息，最大限度地减少损失并组织抗击突发性事件。
星箭一体化飞行器为实现最大的效能，要求具备轻巧的结构，可通过降低运载火箭所运输载荷的重量，例如一体化飞行器中的火箭与卫星，通过共用部分重复的设备和结构，大幅度降低了入轨的冗余重量。但现有星箭一体化中的卫星，其入轨后卫星舱体所对应的载荷比较低，影响了飞行器起飞重量的进一步降低。目前对星箭一体化飞行器系统已经开展了相关技术研究，但是还存在一些问题，具体为：
（1）受制于发射平台承载重量的限制，小型运载器的运载能力有限，使得小型运载器与卫星组合的应用受到限制；
（2）运载器与卫星相互独立导致发射准备过程中组装、测试周期较长，不利于卫星的快速入轨；
（3）运载器与卫星部分设备重复，浪费燃料与硬件资源，存在进一步改善的空间；
（4）入轨飞行器的有效载荷重量占比低；
（5）飞行器的有效载荷、推进系统及电子系统模块化程度低，不利用批量生产和测试。
技术内容
本公开的实施例期望提供一种星箭载一体化飞行器；能够提升运载器的运载能力，提高入轨卫星的有效载荷重量占比的同时，提高了飞行器系统的模块化程度，以实现批量生产和测试。
本公开的实施例的技术方案是这样实现的：
本公开的实施例提供了一种星箭载一体化飞行器，所述星箭载一体化飞行器包括卫星和运载器，所述卫星包括有效载荷、仪器舱、动力控制系统及安装于所述动力控制系统中的贮箱外部上的太阳能帆板，所述运载器包括运载子级一级火箭，运载子级二级火箭，运载子级三级火箭和运载子级末级火箭；其中，
所述卫星与所述运载器的运载子级末级火箭共用所述动力控制系统中的贮箱以形成星箭一体化飞行器；
所述星箭一体化飞行器被倒置地安装于所述运载子级三级火箭与所述运载子级末级火箭之间的过渡段上，以通过所述有效载荷与所述运载器进行连接。
本公开的实施例提供了一种星箭载一体化飞行器；通过卫星与运载器的运载子级末级火箭共用推进系统以形成星箭一体化飞行器，进而将星箭一体化飞行器安装于过渡段上，以实现卫星、运载器以及有效载荷的一体化设计，并同时通过控制有效载荷的重量占比，以提高卫星入轨时的有效载荷的重量占比，进而使得星箭载一体化飞行器能够更好地实现对突发性事件的观测。
 附图简要说明
为了更清楚地说明本公开实施例的技术方案，下面将对实施例的描述中所需要使用的附图作简单的介绍。下面描述中的附图仅仅是本公开的示例性实施例。
图1为本公开实施例提供的一种星箭载一体化飞行器的结构示意图；
图2为图1中A部分的局部放大图；
图3为本公开实施例提供的卫星结构的爆炸示意图；
图4为常规技术方案中卫星与运载子级末级火箭之间的连接示意图；
图5为本公开实施例提供的有效载荷的主结构与仪器舱的安装板之间的连接示意图；
图6为图5中沿B‑B的截面示意图；
图7为本公开实施例提供的有效载荷与仪器舱的结构组装示意图；
图8为本公开实施例提供的有效载荷与仪器舱之间的传热示意图；
图9为本公开实施例提供的卫星的结构组装示意图；
图10为本公开实施例提供的动力控制系统的结构示意图。
 具体实施方式
为了使得本公开的目的、技术方案和优点更为明显，下面将参照附图详细描述根据本公开的示例实施例。显然，所描述的实施例仅仅是本公开的一部分实施例，而不是本公开的全部实施例，应理解，本公开不受这里描述的示例实施例的限制。
下面将结合本公开实施例中的附图，对本公开实施例中的技术方案进行清楚、完整地描述。
参见图1至图3，其示出了本公开实施例提供的一种星箭载一体化飞行器1，所述星箭载一体化飞行器1包括卫星11和运载器12，所述卫星11包括有效载荷111、仪器舱112、动力控制系统113及安装于所述动力控制系统113上的太阳能帆板114，所述运载器12包括运载子级一级火箭121，运载子级二级火箭122，运载子级三级火箭123和运载子级末级火箭124；其中，
所述卫星11与所述运载器12的运载子级末级火箭124共用所述动力控制系统113中的贮箱1131以形成星箭一体化飞行器13；
所述星箭一体化飞行器13被倒置地安装于所述运载子级三级火箭123与所述运载子级末级火箭124之间的过渡段1231上，以通过所述有效载荷111与所述运载器12进行连接。
需要说明的是，在常规技术方案中，如图4所示，卫星11与运载子级末级火箭124分别在卫星推力器41与末级推力器42的作用下使得对接环43与运载子级末级火箭124上的卫星支撑舱44分离，进而使得卫星11与运载器12的运载子级末级火箭124分开以形成使得运载器12和卫星11的分体式发射方式，但是，在本公开实施例中运载器12的运载子级末级火箭124与卫星11共用推进系统形成了运载器12与卫星11的一体化发射方式。
对于星箭载一体化飞行器1，在一些可能的实施方式中，如图3所示，所述有效载荷111中设置有柱体的主结构1111，所述主结构1111的上端部外周设置有多个星箭连接点1112，用于连接所述星箭一体化飞行器13与所述过渡段1231。
对于上述实施方式，在一些示例中，所述多个星箭连接点1112与所述过渡段1231之间的连接处分别设置有图5中所示的隔振器52，以减少运载器12的主动段振动对有效载荷111的影响。
对于上述实施方式，在一些示例中，如图5所示，所述主结构1111与所述仪器舱112的安装板31之间通过螺栓相连接，且所述主结构1111与所述仪器舱12的安装板31之间的连接处设置有隔振器52，以降低仪器舱112的振动对有效载荷111的影响。
当然，如图6所示，在具体实施过程中在螺栓51与隔振器52之间还设置有限位衬套61。
可以理解地，如图7所示，仪器舱112通过安装板31与有效载荷111安装后，实现了星载一体化。
此外，如图3所示，仪器舱112用于安装飞行器仪器32，包括电子设备，电源控制器及蓄电池等的电子仪器，以及卫星支撑舱44。
对于上述实施方式，在一些示例中，如图8所示，所述主结构1111与所述仪器舱112的安装板31之间设置有多层导热膜81。可以理解地，在具体实施过程中可以根据不同的热控需求以调节导热膜81的层数来实现不同的传热或隔热效果。
对于上述实施方式，在一些示例中，如图3所示，所述主结构1111与所述动力控制系统113中的贮箱1131之间通过螺栓相连接，且所述主结构1111与所述动力控制系统113中的贮箱1131之间的连接设置有隔振器52，以降低动力控制系统113振动对有效载荷111的影响。
可以理解地，在具体实施过程中，如图3所示，安装板31上设置有安装孔311，以使贮箱1131的底座11311能够通过上述的安装孔311以与有效载荷111的主结构1111通过螺栓连接，当然，在贮箱1131的底座11311上设置有螺栓孔11312。
此外，参见图9，其示出了卫星11组装后的结构示意图。由图9可以看出，运载器12通过有效载荷111上的星箭连接点1112直接传力至卫星11，实现了星箭载一体化飞行器1的整体传力路径最短。另一方面，在本公开实施例中，将运载器12与卫星11一体化设计，实现了星箭载一体化飞行器1的轻量化设计，并提高了入轨卫星11的有效载荷111的重量占比，具体来说，卫星11中其余系统的重量占比为30%，有效载荷111的重量占比可达到70%。
需要说明的是，在本公开实施例中，有效载荷111的主结构1111为主要承力结构，因此在具体实施过程中进行了加强设计，以更好地进行传力。
当然，在本公开实施例中，所述动力控制系统113、所述仪器舱112与所述有效载荷111之间能够独立拆装及测试。可以理解地，卫星11中的有效载荷111、仪器舱112及动力控制系统113各自之间进行模块化设计，能够便于独立组装及测试。具体来说，有效载荷111、仪器舱112及动力控制系统113的模块化设计，实现了快速批量生产、独立批量测试，使得总装和测试时间能够缩短70%。
对于星箭载一体化飞行器1，在一些可能的实施方式中，所述动力控制系统112的贮箱1131与所述仪器舱112的安装板31之间设置有隔振器52，以减少所述贮箱1131中的液体晃动对所述有效载荷111的影响。
对于星箭载一体化飞行器1，在一些可能的实施方式中，如图10所示，所述动力控制系统113的贮箱1131的外部还安装有飞轮101、磁力矩器102、轨控发动机103、多个姿控发动机104及多个高压气瓶105；其中，所述高压气瓶105对称设置于所述贮箱1131的底面上；所述轨控发动机103设置于所述高压气瓶105之间并向外喷射气体产生推进力；所述姿控发动机104对称设置于所述高压气瓶105的外侧并向外喷射气体产生推进力。
对于上述实施方式，在一些示例中，如图10所示，所述贮箱1131与所述飞轮101之间设置有隔振器52，以降低由于飞轮101微振动对有效载荷111的影响。
可以理解地，在本公开实施例中，隔振器52可以为T型橡胶隔振器，也可以根据实际的需求采用其他类型的橡胶隔振器；当然，隔振器52可以根据不同的应用场景采用不同的尺寸。
对于本公开实施例提供的飞行器系统1，其通过将仪器舱112安装于有效载荷111结构上，实现了星载一体化，同时将卫星11与运载器12的运载子级末级火箭124共用推进系统形成星箭一体化飞行器13，进而将星箭一体化飞行器13安装于运载子级三级火箭123与运载子级末级火箭124之间的过渡段1231上，以实现卫星11、运载器12以及有效载荷111的一体化设计，并同时通过控制有效载荷111的重量占比，以提高卫星11入轨时的有效载荷111的重量占比，进而使得星箭载一体化飞行器1能够更好地实现对突发性事件的观测。
在上面详细描述的本公开的示例实施例仅仅是说明性的，而不是限制性的。本领域技术人员应该理解，在不脱离本公开的原理和精神的情况下，可对这些实施例或其特征进行各种修改和组合，这样的修改应落入本公开的范围内。
 工业实用性
本公开的实施例提供了一种星箭载一体化飞行器；通过卫星与运载器的运载子级末级火箭共用推进系统以形成星箭一体化飞行器，进而将星箭一体化飞行器安装于过渡段上，以实现卫星、运载器以及有效载荷的一体化设计，并同时通过控制有效载荷的重量占比，以提高卫星入轨时的有效载荷的重量占比，进而使得星箭载一体化飞行器能够更好地实现对突发性事件的观测。

Claims (10)

  1. 一种星箭载一体化飞行器，其特征在于，所述星箭载一体化飞行器包括卫星和运载器，所述卫星包括有效载荷、仪器舱、动力控制系统及安装于所述动力控制系统中的贮箱外部上的太阳能帆板，所述运载器包括运载子级一级火箭，运载子级二级火箭，运载子级三级火箭和运载子级末级火箭；其中，
    所述卫星与所述运载器的运载子级末级火箭共用所述动力控制系统中的贮箱以形成星箭一体化飞行器；
    所述星箭一体化飞行器被倒置地安装于所述运载子级三级火箭与所述运载子级末级火箭之间的过渡段上，以通过所述有效载荷与所述运载器进行连接。
  2. 根据权利要求1所述的星箭载一体化飞行器，其特征在于，所述有效载荷中设置有柱体的主结构，所述主结构的上端部外周设置有多个星箭连接点，用于连接所述星箭一体化飞行器与所述过渡段。
  3. 根据权利要求2所述的星箭载一体化飞行器，其特征在于，所述多个星箭连接点与所述过渡段之间的连接处分别设置有隔振器。
  4. 根据权利要求2所述的星箭载一体化飞行器，其特征在于，所述主结构与所述仪器舱的安装板之间通过螺栓相连接，且所述主结构与所述仪器舱的安装板之间的连接处设置有隔振器。
  5. 根据权利要求2所述的星箭载一体化飞行器，其特征在于，所述主结构与所述仪器舱的安装板之间设置有多层导热膜。
  6. 根据权利要求2所述的星箭载一体化飞行器，其特征在于，所述主结构与所述动力控制系统中的贮箱之间通过螺栓相连接，且所述主结构与所述动力控制系统中的贮箱之间的连接设置有隔振器。
  7. 根据权利要求1所述的星箭载一体化飞行器，其特征在于，所述动力控制系统、所述仪器舱与所述有效载荷之间能够独立拆装及测试。
  8. 根据权利要求1所述的星箭载一体化飞行器，其特征在于，所述动力控制系统的贮箱与所述仪器舱的安装板之间设置有隔振器。
  9. 根据权利要求1所述的星箭载一体化飞行器，其特征在于，所述动力控制系统的贮箱的外部还安装有飞轮、磁力矩器、轨控发动机、多个姿控发动机及多个高压气瓶；其中，所述高压气瓶对称设置于所述贮箱的底面上；所述轨控发动机设置于所述高压气瓶之间并向外喷射气体产生推进力；所述姿控发动机对称设置于所述高压气瓶的外侧并向外喷射气体产生推进力。
  10. 根据权利要求9所述的星箭载一体化飞行器，其特征在于，所述贮箱与所述飞轮之间设置有隔振器。
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