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WO2017018903A1 - Procédé de mise en orbite de charge utile à l'aide d'un lanceur - Google Patents

Procédé de mise en orbite de charge utile à l'aide d'un lanceur Download PDF

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Publication number
WO2017018903A1
WO2017018903A1 PCT/RU2015/000473 RU2015000473W WO2017018903A1 WO 2017018903 A1 WO2017018903 A1 WO 2017018903A1 RU 2015000473 W RU2015000473 W RU 2015000473W WO 2017018903 A1 WO2017018903 A1 WO 2017018903A1
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Prior art keywords
rocket
central
thrust
engine
engines
Prior art date
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Ceased
Application number
PCT/RU2015/000473
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English (en)
Russian (ru)
Inventor
Сергей Борисович БЫКОВСКИЙ
Павел Сергеевич ПУШКИН
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Cosmocourse (llc Cosmocourse) LLC
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Cosmocourse (llc Cosmocourse) LLC
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Priority to PCT/RU2015/000473 priority Critical patent/WO2017018903A1/fr
Publication of WO2017018903A1 publication Critical patent/WO2017018903A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles

Definitions

  • the invention relates to the field of rocket and space technology and can find application in the creation and modernization of missile systems for various purposes, including means of launching payloads into low Earth orbit and launching payloads to a suborbital ballistic trajectory.
  • the known method includes attaching, in accordance with the launch program, to the central missile unit of a tandemly located booster missile unit and a head unit with a payload, forming a lower multi-block package of missile units by attaching side missile units to the central missile unit, turning on the start of all marching rocket engine side and central missile blocks, the joint operation of the main and lateral rocket blocks marching liquid propellant rocket engines until the side rocket blocks generate fuel, turning off the neck rocket engine of the side rocket blocks and separation of the side rocket blocks from the central rocket block while continuing to operate the mid-range rocket engine of the central rocket block before fuel is generated from it, shutdown of the main rocket rocket engine of the central rocket block, separation of the tandem located booster rocket and head block from the central rocket block, and the subsequent overclocking the head unit until it goes into orbit.
  • the disadvantage of this scheme is that the central missile unit has large dimensions and mass compared to the side missile units and carries more fuel in its tanks, which ensures a longer operation of its marching rocket engine. This affects the energy capabilities of the packet scheme, since the mass of the central unit includes a large fraction of the mass of the structure that provides fuel storage for the operation of the rocket engine at the stage of flight of the first stage.
  • the closest analogue is the solution "method of putting into payload the orbit of a multifunctional launch vehicle of a combined circuit with marching liquid propellant rocket propulsion systems (jets), a multifunctional launch rocket of a combined circuit with marching rocket engines and a method for its development" described in RF patent jN ° 2161 108, published December 27, 2000.
  • the patent discloses Angara launch vehicles with a packet arrangement of rocket blocks of the first and second stages.
  • a combined circuit is used with a lower multiblock package of identical rocket blocks having adjustable marching rocket engines with the same nominal thrust, when the launch vehicle starts, the marching rocket engines of the side rocket blocks are brought to rated thrust, and the main marching rocket engine of the central missile block - to thrust equal to 90 ... 100% of the nominal value, and maintain it unchanged until the launch vehicle reaches longitudinal acceleration of 12.7 ... 16.7 m / s (1.3 ... 1.7 g), then reduce the thrust of the mid-range rocket engine missile block up to 0.3 ...
  • the thrust of the marching rocket engine of the central unit is increased to the nominal value.
  • the use of adjustable marching liquid propellant rocket engines in these missile blocks allows to fully realize the energy capabilities of the lower multiblock package of missile blocks at the start, and a subsequent decrease in the thrust of the main missile propellant rocket engine to 0.3 ... 0.5 of the nominal thrust ensures that it remains in the central rocket fuel supply unit for its marching rocket engine after separation of the side rocket blocks.
  • the thrust reduction of the marching liquid propellant rocket engine of the central missile unit begins after the launch vehicle reaches a longitudinal acceleration of 12.7 ... 16.7 m / s2 (1.3 ...
  • the problem to which this invention is directed is to create a method of flight of launch vehicles with a packet arrangement of rocket blocks of the first and second stages, which allows to increase the mass of the payload.
  • the technical result is to increase the load capacity of the exploited and developed launch vehicles with minimal changes in their design.
  • a method for putting payload into orbit with a launch vehicle with a multiblock package of missile units of a combined scheme includes the following steps: a. at the launch of the launch vehicle, marching liquid propellant propulsion systems (LRE) of the lateral and central missile units are launched for nominal thrust,
  • LRE liquid propellant propulsion systems
  • At least one engine of the mid-range main propellant rocket engine is switched off or throttled to a level below 0.3 of the nominal thrust
  • the head unit is brought out, including the tandemly arranged upper steps to a predetermined path.
  • stage b. occurs when the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s.
  • they can form a lower multiblock package of missile blocks with non-identical fuel tanks, mass-dimensional characteristics and mid-range rocket engines with the same or different nominal thrust. They can also form a lower multiblock package of rocket blocks with various components of rocket fuel.
  • the thrust of the marching liquid propellant rocket engine of the central missile unit can be reached and maintained unchanged to a value that ensures that the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s 2 .
  • the inclusion of liquid propellant rocket engines can be performed before the marching liquid propellant rocket engines of the side rocket blocks are turned off, sufficient for the central rocket engine to reach the nominal operating mode.
  • they can install at least one marching propulsion engine rocket engine of the central rocket block, which is turned off or throttled to a level less than 0.3 of the nominal thrust, to move the nozzle nozzles, which are shifted after turning off or throttling the engine to a level less than 0.3 from rated traction.
  • they can install at least two engines of a mid-range main propellant rocket engine, which shut off or throttle to a level less than 0.3 of the nominal thrust, install a single movable nozzle nozzle, which is shifted after shutting down or throttling the engines to a level less than 0, 3 from rated traction.
  • FIG. 1 three-stage launch vehicle with a multiblock package of five identical missile units.
  • FIG. 2 two-stage launch vehicle with a multiblock package of five identical missile units.
  • FIG. 4-PH with a movable nozzle nozzle throughout the LRE of the central unit (a single nozzle nozzle).
  • FIG. 5 three-stage launch vehicle with a multi-block package of three missile units with an excellent rocket engine on the central missile unit.
  • FIG. 6 three-stage launch vehicle with a multiblock package of three identical missile units.
  • FIG. 7 two-stage launch vehicle with a multiblock package of three identical missile units.
  • FIG. 8 is a flight diagram of a three-stage launch vehicle with a central propellant rocket engine shut down and its subsequent inclusion before separation of the side missile blocks
  • FIG. 9 is a flight diagram of a three-stage launch vehicle with a central rocket engine rocket engine shut down and then only part of its rocket engine turned on before separation of the side rocket blocks.
  • pos. 1 central missile unit pos. 2 - lateral missile unit pos. 3 - transition compartment pos. 4 - missile unit of the third stage pos. 5 - head unit pos.6 - LRE rocket of the side rocket block pos.7 - LRS of the central missile block pos.8 - nozzle nozzles for one LRE pos.9 - nozzle single nozzle for the whole LRE of pos.10 - shutter of the head fairing pos.11 - payload
  • LRE liquid propellant propulsion systems
  • the liquid propellant rocket engine consists of several single liquid propellant rocket engines (the first stage of the Falcon rocket), then not all liquid propellant rockets can be turned off, but only a part of them. This is similar to the throttling of the LRE of the central block of the Angara launch vehicle.
  • each individual engine has the ability to autonomously control traction and turn on / off, which is implemented in almost all launch vehicles. When the entire rocket engine shuts down, then commands are issued to all engines at once.
  • engines such as 11D58 (booster block type “DM”), S.98M (booster block type “breeze”) and other engines of booster blocks have the ability to re-enable in flight.
  • the Merlin-ID engine on the Falcon-9 LV has the ability to re-enable the flight, which it implements when launching the spent missile unit.
  • Electric ignition may be used.
  • One of its varieties is laser ignition (patent application of the Russian Federation No. 2012157504 company Spectralazer, published on 10.07.2014). In this case, it may be necessary to refine the design itself for re-inclusion.
  • the rocket engine of the central missile unit should be turned off during the launch of the rocket into orbit, but a positive effect, though to a lesser extent, will be in the case of a decrease in the thrust of the main rocket engine central missile unit to a non-zero level, but below 0.3, i.e. in case of throttling to a level below 0.3.
  • the liquid propellant rocket shutdown can occur somewhat earlier or later than the moment the launch vehicle reaches acceleration 11, 8 ... 16.7. These values can vary significantly and depend on the specific rocket.
  • the marching rocket engines of the side missile blocks are turned off, the marching rocket engines of the central missile block are switched on again or throttled to a level above 0.3.
  • the lateral rocket blocks are separated and dropped, with the central rocket block rocket engine turned on, and after the end of the working fuel reserves in the central rocket block, it is separated from the rocket block with the head block with the subsequent dispersal of the head block before it enters a predetermined orbit.
  • the method can form a lower multiblock package of missile blocks with non-identical fuel tanks, mass-dimensional characteristics and mid-range rocket engines with the same or different nominal thrust. They can also form a lower multiblock package of rocket blocks with various components of rocket fuel. To implement the method, they can form a lower multiblock package of missile blocks with a different number of LRE identical LRE with the same nominal thrust.
  • the thrust of the marching liquid propellant rocket engine of the central missile unit can be reached and maintained unchanged to a value that ensures that the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s.
  • the specific thrust value is found by calculation or experimentally for a specific rocket model, which provides the maximum payload mass and design restrictions.
  • liquid propellant rocket engines can be performed before the marching liquid propellant rocket engines of the side rocket units are disconnected for a time sufficient for the central rocket engine to reach the nominal operating mode. This time is determined by the dynamic characteristics of each particular engine. The range can theoretically be any, although from the conditions of efficiency it can be from 0 to 100 seconds.
  • the inclusion of liquid propellant rocket engines can be performed after separation of the side rocket blocks, i.e. with the engines off. This may be due to the fact that when separating missile blocks, requirements for reducing loads, including aerodynamic ones, may arise. That is, in order to reduce the load during separation, it is possible to share with the turned off engines of the central unit.
  • At least one engine marching liquid propellant rocket engine of the central missile unit which is turned off or throttled to a level less than 0.3, can install a movable nozzle nozzle, which is shifted after turning off or throttling the engine to a level of less than 0.3.
  • Nozzle nozzles can further increase the mass of the payload.
  • the nozzle nozzles on the engine shifted at the start, do not allow the rocket to exit the launch pad, and it is not possible to shift the nozzles during the flight on a running engine due to gas dynamics problems.
  • the proposed solution in the case of the engine turned off or throttling close to zero, it turns out to shift the nozzles during the flight.
  • At least two engines of a mid-range main propellant rocket engine which shut off or throttle to a level of less than 0.3, can install a single movable nozzle nozzle, which is shifted after turning off or throttling the engines to a level of less than 0, 3.
  • the engines of the central block After completion of the flight stage of the 1st stage and separation of the side blocks, the engines of the central block are again brought back to the nominal mode.
  • t is the total operating time of all engines of the 1st stage.
  • the characteristic speed that the launch vehicle picks up on the flight section of the 1st and 2nd stages is the characteristic speed that the launch vehicle picks up on the flight section of the 1st and 2nd stages:
  • the payload mass of the 3rd stage M mon which is the payload of the entire launch vehicle, is associated with the initial mass of the 3rd stage M 03 by the ratio:
  • M mon ⁇ ⁇ 3 - M 03.
  • ⁇ advise ⁇ 3 is the relative mass of the 3rd stage payload, which depends on the characteristic speed developed in the flight section of the 3rd stage and the mass perfection of the 3rd stage design.
  • the trajectory parameters do not change; therefore, we consider the value of the characteristic velocity unchanged.
  • Mass excellence structures while increasing the absolute mass of the structure only improves, so the ability to increase the starting mass of the 3rd stage by 5.4% will increase the payload mass by at least 5.4%, which will be 1 on the scale of the Angara-5 launch vehicle , 3 t.
  • the nozzle nozzles were removed to provide maximum earth thrust for the engines of the central unit.
  • the specific thrust impulse of engines with a nozzle nozzle in a void can range from 343.5 to 358.7 kgf ⁇ s / kg (see table 1).
  • Equation (1) we substitute these values into equation (1) as P yD 2. Then, solving equation (1) by numerical methods, we obtain the following expected range of increase in the initial mass of the 3rd stage when the central unit engines are turned off together with the use of nozzle nozzles in the flight section of the 2nd stage:
  • the use of movable nozzle nozzles in the flight section of the 2nd stage can, in addition to turning off the engines of the central block in the flight section of the 1st stage ( ⁇ TM, ⁇ 5.4%), increase in the payload of 1.9 ... 6 ,8 %.
  • Shprp the mass of the control system and all other compartments of the side blocks of the 1st stage.
  • the total mass of fuel of the side and central blocks consumed throughout the flight section of the 1st stage M t1 can be expressed through its starting mass:
  • ⁇ 12 ⁇ ⁇ ⁇ ⁇ (7)
  • k d is the current throttle coefficient of the engines of the central unit.
  • M t12 M t1 - P - / s d cf (8) where k d cf is the average integral throttle coefficient of the engines of the central unit:
  • ⁇ is the mass coefficient of the fuel compartments of the side blocks of the 1st stage (the ratio of the mass of the fuel compartments to the mass of fuel in these compartments).
  • n 01 is the starting overload of the 1st stage.
  • ⁇ ⁇ ⁇ M 0 1 - M 01 (1 - ⁇ ⁇ ) - M 01 (1 - ⁇ ⁇ ) ( ⁇ - P ⁇ c d av ) ⁇ a d - ⁇ 0 ⁇ ⁇ ⁇ ( ⁇ ⁇ P) ⁇ ⁇ ⁇ ⁇ ⁇
  • M ' M H2 02 - m2 - mn TO2 - dB2 t - t np2 (19) wherein M 02 - initial weight of 2nd stage,
  • M is the mass of fuel T2 of the central unit, expended in the area of flight stage 2,
  • the mass of fuel of the central unit consumed in the flight section of the 2nd stage M t2 can be expressed in terms of the initial mass of the 2nd stage:
  • ⁇ ⁇ 2 ⁇ 02 ⁇ (1 - ⁇ ⁇ 2) (20) where ⁇ 2 is the relative final mass of the second stage.
  • the mass of the fuel compartment of the central unit is written taking into account the fact that part of the fuel of the central unit is also consumed in the flight section of the 1st stage. Then, taking into account expressions (20), (8) and (4):
  • y2 is the mass coefficient of the engines of the central unit (the ratio of the mass of the engines of the 2nd stage to the starting thrust of these engines P 012 ) .
  • ⁇ 2 is the mass coefficient of the control system and all other compartments (the ratio of the mass of the control system and all other compartments of the central unit to the initial mass of the 2nd stage).
  • M PN 2 M 02 - M 02 (1 - ⁇ ⁇ 2 ) - M 02 (1 - ⁇ 2 ) 2 - Af 01 (l - those.:
  • the tanks of the central unit should be 2 ... 3 times lighter than the tanks of the side unit, which is practically not feasible.

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  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Testing Of Engines (AREA)

Abstract

L'invention se rapporte au domaine des fusées spatiales et peut être utilisée pour la création et la modernisation de systèmes de fusée à vocations diverses, y compris pour mettre en orbite basse des charges utiles et placer des charges utiles en trajectoire balistique suborbitale. L'invention concerne essentiellement un procédé de mise en orbite de charge utile à l'aide d'un lanceur comprenant un ensemble à unités multiples d'unités de fusées d'un circuit combiné, lequel procédé consiste à: a) lors l'allumage du lanceur on amène les moteurs à propergol liquide (MPL) principaux des unités de fusées latérales et centrale à la poussée nominale; b) une fois que le lanceur a atteint un accélération longitudinale assurant son positionnement stable sur la trajectoire, on éteint au moins moteur MPL principal de l'unité de fusée centrale ou on le ralentit jusqu'à un niveau représentant moins de 0,3 fois la poussée nominale; c) avant l'extinction des LP principaux des unités de fusées latérales, on procède un allumage ou un ajustement jusqu'à un niveau représentant plus de 0,3 fois la poussée nominale du moteur MLP de l'unité de fusée centrale dont la poussée avait été diminuée; d) on sépare et on éjecte les unités de fusées latérales lorsque le MPL de l'unité de fusée centrale est allumé; e) on place l'étage supérieur en orbite. Le résultat technique consiste en une plus grande capacité d'emport de charge des lanceurs en exploitation ou à produire tout en respectant des modifications minimales de leur structure.
PCT/RU2015/000473 2015-07-28 2015-07-28 Procédé de mise en orbite de charge utile à l'aide d'un lanceur Ceased WO2017018903A1 (fr)

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Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN109711010A (zh) * 2018-12-13 2019-05-03 北京航天自动控制研究所 一种垂直起降火箭在线轨迹规划的发动机特性处理方法
CN112596537A (zh) * 2020-11-27 2021-04-02 中国人民解放军国防科技大学 用于在线轨迹规划的模型误差补偿方法、系统及存储介质
CN112966340A (zh) * 2021-01-19 2021-06-15 中国人民解放军63921部队 一种固体捆绑运载火箭运载能力的小偏差快速修正方法
CN113212808A (zh) * 2021-05-08 2021-08-06 北京格锐德科技有限公司 一种基于挤压发动机的运载火箭
CN113642097A (zh) * 2021-06-28 2021-11-12 上海宇航系统工程研究所 一种基于统计能量法的捆绑火箭高频环境仿真预示方法
CN114018103A (zh) * 2021-11-08 2022-02-08 航天科工火箭技术有限公司 一种基于小推力的运载火箭弹道重构方法及系统
CN114060171A (zh) * 2021-09-14 2022-02-18 航天科工火箭技术有限公司 一种火箭以及火箭推进剂晃动抑制方法和装置
CN114216376A (zh) * 2021-12-09 2022-03-22 北京航天自动控制研究所 运载火箭的多载荷分级优化方法

Citations (3)

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US5141181A (en) * 1989-10-05 1992-08-25 Leonard Byron P Launch vehicle with interstage propellant manifolding
EP0508609B1 (fr) * 1991-04-08 1998-05-20 Trw Inc. Lanceur modulaire à propergol solide et système de lancement
RU2161108C1 (ru) * 2000-02-07 2000-12-27 Государственный космический научно-производственный центр им. М.В. Хруничева Способ выведения на орбиту полезной нагрузки многофункциональной ракетой-носителем комбинированной схемы с маршевыми жидкостными ракетными двигательными установками (жрду), многофункциональная ракета-носитель комбинированной схемы с маршевыми жрду и способ ее отработки

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5141181A (en) * 1989-10-05 1992-08-25 Leonard Byron P Launch vehicle with interstage propellant manifolding
EP0508609B1 (fr) * 1991-04-08 1998-05-20 Trw Inc. Lanceur modulaire à propergol solide et système de lancement
RU2161108C1 (ru) * 2000-02-07 2000-12-27 Государственный космический научно-производственный центр им. М.В. Хруничева Способ выведения на орбиту полезной нагрузки многофункциональной ракетой-носителем комбинированной схемы с маршевыми жидкостными ракетными двигательными установками (жрду), многофункциональная ракета-носитель комбинированной схемы с маршевыми жрду и способ ее отработки

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN109711010A (zh) * 2018-12-13 2019-05-03 北京航天自动控制研究所 一种垂直起降火箭在线轨迹规划的发动机特性处理方法
CN109711010B (zh) * 2018-12-13 2023-05-12 北京航天自动控制研究所 一种垂直起降火箭在线轨迹规划的发动机特性处理方法
CN112596537B (zh) * 2020-11-27 2022-03-29 中国人民解放军国防科技大学 用于在线轨迹规划的模型误差补偿方法、系统及存储介质
CN112596537A (zh) * 2020-11-27 2021-04-02 中国人民解放军国防科技大学 用于在线轨迹规划的模型误差补偿方法、系统及存储介质
CN112966340A (zh) * 2021-01-19 2021-06-15 中国人民解放军63921部队 一种固体捆绑运载火箭运载能力的小偏差快速修正方法
CN112966340B (zh) * 2021-01-19 2023-12-29 中国人民解放军63921部队 一种固体捆绑运载火箭运载能力的小偏差快速修正方法
CN113212808A (zh) * 2021-05-08 2021-08-06 北京格锐德科技有限公司 一种基于挤压发动机的运载火箭
CN113642097B (zh) * 2021-06-28 2023-07-14 上海宇航系统工程研究所 一种基于统计能量法的捆绑火箭高频环境仿真预示方法
CN113642097A (zh) * 2021-06-28 2021-11-12 上海宇航系统工程研究所 一种基于统计能量法的捆绑火箭高频环境仿真预示方法
CN114060171A (zh) * 2021-09-14 2022-02-18 航天科工火箭技术有限公司 一种火箭以及火箭推进剂晃动抑制方法和装置
CN114018103A (zh) * 2021-11-08 2022-02-08 航天科工火箭技术有限公司 一种基于小推力的运载火箭弹道重构方法及系统
CN114216376A (zh) * 2021-12-09 2022-03-22 北京航天自动控制研究所 运载火箭的多载荷分级优化方法
CN114216376B (zh) * 2021-12-09 2023-11-14 北京航天自动控制研究所 运载火箭的多载荷分级优化方法

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