JPH06280680A - Method and apparatus for controlling thrust of two-liquid system rocket engine - Google Patents
Method and apparatus for controlling thrust of two-liquid system rocket engineInfo
- Publication number
- JPH06280680A JPH06280680A JP6714393A JP6714393A JPH06280680A JP H06280680 A JPH06280680 A JP H06280680A JP 6714393 A JP6714393 A JP 6714393A JP 6714393 A JP6714393 A JP 6714393A JP H06280680 A JPH06280680 A JP H06280680A
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- Prior art keywords
- fuel
- oxidant
- combustor
- valve
- thrust
- Prior art date
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Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は宇宙空間で使用するロケ
ットエンジンに係わり、更に詳しくは二液式ロケットエ
ンジンに関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a rocket engine used in outer space, and more particularly to a two-component rocket engine.
【０００２】[0002]
【従来の技術】宇宙空間で使用するロケットエンジンに
は、一液式エンジンと二液式エンジンとがある。一液式
エンジンは、燃料を触媒で自己分解させ、分解した燃料
をノズルを通して噴射するものであり、構造が簡単であ
るが、噴射圧力が低く、比推力が小さい（例えば約２２
０ｓｅｃ）、等の問題点がある。これに対し、二液式エ
ンジンは、燃料と酸化剤を燃焼室内に別々に供給し、燃
焼室内で燃料を燃焼させ、燃焼ガスを噴射ノズルを通し
て外部に噴射し、推力を得るものであり、噴射圧力が高
く、比推力が大きい（例えば約３００ｓｅｃ）、等の特
徴があり、大推力を必要とする場合に広く用いられてい
る。Rocket engines used in outer space include one-component engines and two-component engines. A one-liquid engine has a simple structure, in which fuel is self-decomposed by a catalyst and the decomposed fuel is injected through a nozzle. However, the injection pressure is low and the specific thrust is small (for example, about 22).
0 sec), etc. On the other hand, a two-component engine supplies fuel and an oxidant separately into the combustion chamber, burns the fuel in the combustion chamber, injects combustion gas to the outside through an injection nozzle, and obtains thrust. It is characterized by high pressure and large specific thrust (for example, about 300 sec), and is widely used when a large thrust is required.
【０００３】[0003]
【発明が解決しようとする課題】かかる二液式ロケット
エンジンの推力制御は、燃料タンク及び酸化剤タンクを
不活性ガスで加圧し、このガス圧力を制御することによ
り、燃焼器に供給する燃料及び酸化剤の流量を制御して
燃焼器から噴射する燃焼ガスの流量及び速度を変化させ
る手段が従来から用いられていた。しかし、かかる手段
では、推力を小さくする（すなわち、ガス圧力を下げて
燃料及び酸化剤の流量を少なくする）と、燃焼器内で振
動燃焼が起こり、燃焼器内の燃焼圧力が大きく変動する
ため、燃料及び酸化剤の流量も大きく変動し、一定の安
定した推力を得ることが難しい問題点があった。かかる
問題点を解決するため、キャビテーションバルブにより
流量制御を行う推力制御手段が提案され一部で使用され
た。しかし、かかるキャビテーションバルブは縮流部が
キャビテーションを起こすキャビテーションオリフィス
を用い、その縮流部面積をニィードル等で変化させるも
のであり、遠隔操作でニィードルを移動させる機構が複
雑である問題点があった。また、大推力を得るためには
縮流部を相当大きくする必要があり、オリフィスの形状
を大推力に適合させるために特殊形状にする必要がある
問題点があった。The thrust control of such a two-component rocket engine is performed by pressurizing the fuel tank and the oxidizer tank with an inert gas, and controlling the gas pressure to control the fuel and the fuel supplied to the combustor. Conventionally, means for controlling the flow rate of the oxidizer to change the flow rate and speed of the combustion gas injected from the combustor has been used. However, in such means, when the thrust is reduced (that is, the gas pressure is reduced to reduce the flow rates of the fuel and the oxidant), oscillatory combustion occurs in the combustor, and the combustion pressure in the combustor fluctuates greatly. However, the flow rates of the fuel and the oxidizer also fluctuate greatly, and it is difficult to obtain a constant and stable thrust. In order to solve such a problem, a thrust control means for controlling the flow rate by a cavitation valve has been proposed and partially used. However, such a cavitation valve uses a cavitation orifice that causes cavitation in the contraction part, and changes the area of the contraction part with a nidle, etc., and there is a problem that the mechanism for moving the nidle by remote control is complicated. . Further, in order to obtain a large thrust, it is necessary to make the contraction portion considerably large, and there is a problem that the shape of the orifice needs to be a special shape in order to adapt to the large thrust.
【０００４】本発明はかかる問題点を解決するために創
案されたものである。すなわち、本発明の目的は、二液
式ロケットエンジンの推力を大推力から小推力まで広範
囲に容易に制御することができる制御方法及び装置を提
供することにある。The present invention was created to solve such problems. That is, it is an object of the present invention to provide a control method and device capable of easily controlling the thrust of a two-component rocket engine in a wide range from a large thrust to a small thrust.
【０００５】[0005]
【課題を解決するための手段】本発明によれば、ガスで
加圧される燃料タンク及び酸化剤タンクと、燃料を酸化
剤で燃焼させる燃焼器と、燃料タンクから燃料を燃焼器
に導く燃料ラインと、酸化剤タンクから酸化剤を燃焼器
に導く酸化剤ラインとを備え、更に、燃料ラインに、燃
料を燃焼器に直接導く主燃料遮断弁と、燃料を燃料用キ
ャビテーションオリフィスを介して燃焼器に導く副燃料
遮断弁とを、互いに並列に設け、酸化剤ラインに、酸化
剤を燃焼器に直接導く主酸化剤遮断弁と、酸化剤を酸化
剤用キャビテーションオリフィスを介して燃焼器に導く
副酸化剤遮断弁とを、互いに並列に設け、大推力時に、
前記主燃料遮断弁と主酸化剤遮断弁を開放し、副燃料遮
断弁と副酸化剤遮断弁を閉鎖し、小推力時に、前記主燃
料遮断弁と主酸化剤遮断弁を閉鎖し、副燃料遮断弁と副
酸化剤遮断弁を開放し、かつ、燃料タンク及び酸化剤タ
ンクを加圧するガス圧を制御して、燃焼器に導びかれる
燃料及び酸化剤の流量を調節する、ことを特徴とする二
液式ロケットエンジンの推力制御方法及び装置が提供さ
れる。According to the present invention, a fuel tank and an oxidizer tank that are pressurized with gas, a combustor that burns the fuel with the oxidizer, and a fuel that guides the fuel from the fuel tank to the combustor Line, and an oxidant line that guides the oxidant from the oxidant tank to the combustor.Furthermore, a main fuel cutoff valve that directly guides the fuel to the combustor and a fuel is burned through the fuel cavitation orifice to the fuel line. A secondary fuel cutoff valve that leads to the burner is provided in parallel with each other, and a main oxidant cutoff valve that directly guides the oxidant to the combustor and an oxidant to the burner through the cavitation orifice for the oxidizer in the oxidizer line. A secondary oxidant shutoff valve is installed in parallel with each other,
The main fuel cutoff valve and the main oxidant cutoff valve are opened, the sub fuel cutoff valve and the sub oxidant cutoff valve are closed, and at the time of a small thrust, the main fuel cutoff valve and the main oxidant cutoff valve are closed, and the sub fuel The shutoff valve and the auxiliary oxidant shutoff valve are opened, and the gas pressure for pressurizing the fuel tank and the oxidant tank is controlled to adjust the flow rates of the fuel and the oxidant introduced to the combustor. A method and apparatus for controlling thrust of a two-component rocket engine are provided.
【０００６】[0006]
【作用】本発明によれば、大推力時には、主燃料遮断弁
と主酸化剤遮断弁が開放されているので、この主燃料遮
断弁及び主酸化剤遮断弁を介して燃料と酸化剤が燃料タ
ンク及び酸化剤タンクから直接燃焼器に導入される。従
って、燃料タンク及び酸化剤タンクを加圧するガス圧を
制御することにより、燃焼器に導びかれる燃料及び酸化
剤の流量を調節することができる。この構成は、従来の
一般的な二液式ロケットエンジンの推力制御と同様であ
り、大推力時に燃焼器に供給する燃料及び酸化剤の流量
を安定して制御し、燃焼器から噴射する燃焼ガスの流量
及び速度を変化させ、大推力を安定して調節することが
できる。一方、この構成のままガス圧を下げて燃料及び
酸化剤の流量を少なくすると、前述のように、燃焼器内
で振動燃焼が起こり、燃焼器内の燃焼圧力が大きく変動
する。しかし、本発明によれば、小推力時に、主燃料遮
断弁と主酸化剤遮断弁が閉鎖され、副燃料遮断弁と副酸
化剤遮断弁が開放されるので、副燃料遮断弁及び副酸化
剤遮断弁を通り、燃料用キャビテーションオリフィス及
び酸化剤用キャビテーションオリフィスを介して燃料と
酸化剤が燃料タンク及び酸化剤タンクから燃焼器にそれ
ぞれ導入される。キャビテーションオリフィスはその縮
流部でキャビテーションが発生しており、その部分の圧
力は燃料及び酸化剤の飽和圧力であり一定である。従っ
て、燃焼器内で振動燃焼が起こり、燃焼器内の燃焼圧力
が大きく変動しても、燃料及び酸化剤の流量は燃焼器内
の燃焼圧力に影響されず上流のガス圧のみで制御するこ
とができ、一定の安定した小推力を得ることができる。According to the present invention, since the main fuel cutoff valve and the main oxidant cutoff valve are opened at the time of a large thrust, the fuel and the oxidant are fed through the main fuel cutoff valve and the main oxidant cutoff valve. It is introduced directly into the combustor from the tank and the oxidant tank. Therefore, by controlling the gas pressure that pressurizes the fuel tank and the oxidant tank, the flow rates of the fuel and the oxidant introduced into the combustor can be adjusted. This structure is similar to the thrust control of a conventional general two-liquid rocket engine, and the flow rate of the fuel and oxidant supplied to the combustor at the time of large thrust is stably controlled, and the combustion gas injected from the combustor is controlled. It is possible to stably adjust the large thrust force by changing the flow rate and speed of. On the other hand, if the gas pressure is reduced and the flow rates of the fuel and the oxidant are reduced with this configuration, as described above, oscillatory combustion occurs in the combustor, and the combustion pressure in the combustor fluctuates greatly. However, according to the present invention, when the thrust is small, the main fuel cutoff valve and the main oxidant cutoff valve are closed, and the sub fuel cutoff valve and the sub oxidant cutoff valve are opened. The fuel and the oxidant are introduced into the combustor from the fuel tank and the oxidant tank through the cutoff valve and the cavitation orifice for the fuel and the cavitation orifice for the oxidant, respectively. In the cavitation orifice, cavitation is generated in its contracted portion, and the pressure in that portion is a saturated pressure of the fuel and the oxidizer and is constant. Therefore, even if oscillatory combustion occurs in the combustor and the combustion pressure in the combustor fluctuates significantly, the flow rates of fuel and oxidant should not be affected by the combustion pressure in the combustor and should be controlled only by the upstream gas pressure. Therefore, a constant and stable small thrust can be obtained.
【０００７】[0007]
【実施例】以下、本発明の好ましい実施例を図面を参照
して説明する。図１は、本発明による二液式ロケットエ
ンジンの推力制御装置の全体構成図である。この図にお
いて、本発明による二液式ロケットエンジンは、液体燃
料を液体酸化剤で燃焼させるエンジンであり、加圧タン
ク１に充填されたガスで加圧される燃料タンク２及び酸
化剤タンク４と、燃料を酸化剤で燃焼させる燃焼器６
と、燃料タンク２から燃料を燃焼器６に導く燃料ライン
８と、酸化剤タンク４から酸化剤を燃焼器６に導く酸化
剤ライン１０とを備えている。加圧タンク１の出口ライ
ンには調圧器３が設けられ、燃料タンク２及び酸化剤タ
ンク４を加圧する圧力を調節するようになっている。な
お、図示のように、燃料タンク２及び酸化剤タンク４を
単一の加圧タンク１と調圧器３で加圧してもよく、又
は、それぞれ独立に加圧タンク及び調圧器を備えてもよ
く、或いは調圧器のみをそれぞれ独立に備えてもよい。
燃料タンク２及び酸化剤タンク４はそれぞれ液体の燃
料、酸化剤を内部に収容しており、ガス圧により外部に
燃料、酸化剤が流出するようになっている。燃料は、ヒ
ドラジン系燃料、例えばヒドラジン（Ｎ2 Ｈ4 ）又はモ
ノメチルヒドラジン（ＭＭＨ）等であり、酸化剤は、例
えば４酸化２窒素（Ｎ2 Ｏ4）である。DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT A preferred embodiment of the present invention will be described below with reference to the drawings. FIG. 1 is an overall configuration diagram of a thrust control device for a two-component rocket engine according to the present invention. In this figure, a two-component rocket engine according to the present invention is an engine that burns liquid fuel with a liquid oxidant, and includes a fuel tank 2 and an oxidant tank 4 that are pressurized with a gas filled in a pressure tank 1. , Combustor 6 that burns fuel with oxidant
And a fuel line 8 for guiding the fuel from the fuel tank 2 to the combustor 6, and an oxidant line 10 for guiding the oxidant from the oxidant tank 4 to the combustor 6. A pressure regulator 3 is provided on the outlet line of the pressure tank 1 so as to adjust the pressure for pressurizing the fuel tank 2 and the oxidant tank 4. As shown in the figure, the fuel tank 2 and the oxidizer tank 4 may be pressurized by the single pressure tank 1 and the pressure regulator 3, or may be independently provided with the pressure tank and the pressure regulator. Alternatively, only the pressure regulators may be independently provided.
Each of the fuel tank 2 and the oxidant tank 4 contains a liquid fuel and an oxidant inside, and the fuel and the oxidant flow out to the outside due to gas pressure. The fuel is a hydrazine-based fuel such as hydrazine (N 2 H 4 ) or monomethylhydrazine (MMH), and the oxidant is, for example, dinitrogen tetraoxide (N 2 O 4 ).
【０００８】本発明による二液式ロケットエンジンの推
力制御装置は、更に燃料ライン８に互いに並列に設けら
れた主燃料遮断弁１２と副燃料遮断弁１３とを備える。
主燃料遮断弁１２は、燃料を燃料タンク２から燃焼器６
に直接導くようになっている。副燃料遮断弁１３の下流
側には燃料用キャビテーションオリフィス１４が設けら
れており、副燃料遮断弁１３により燃料用キャビテーシ
ョンオリフィス１４を介して燃料を燃焼器６に導くよう
になっている。燃料用キャビテーションオリフィス１４
は、縮流部がキャビテーションを起こす寸法に形成され
ており、キャビテーションを起こすと縮流部の圧力が燃
料の蒸発圧力となり、下流側の圧力に影響されずに上流
側の圧力のみで流れる流量を調節することができる。同
様にこの装置は、酸化剤ライン１０に互いに並列に設け
られた、酸化剤を燃焼器６に直接導く主酸化剤遮断弁１
５と、酸化剤を燃焼器６に酸化剤用キャビテーションオ
リフィス１７を介して導く副酸化剤遮断弁１６とを備え
ている。The thrust control system for a two-pack rocket engine according to the present invention further comprises a main fuel cutoff valve 12 and an auxiliary fuel cutoff valve 13 which are provided in parallel in the fuel line 8.
The main fuel cutoff valve 12 transfers the fuel from the fuel tank 2 to the combustor 6
It is designed to lead you directly to. A cavitation orifice 14 for fuel is provided on the downstream side of the sub fuel cutoff valve 13, and the fuel is guided to the combustor 6 via the cavitation orifice 14 for fuel by the sub fuel cutoff valve 13. Cavitation orifice 14 for fuel
Is formed in such a size that the contraction part causes cavitation, and when cavitation occurs, the pressure in the contraction part becomes the evaporation pressure of the fuel, and the flow rate that flows only by the upstream pressure is not affected by the downstream pressure. It can be adjusted. Similarly, this device comprises a main oxidant shut-off valve 1 provided in parallel with each other in an oxidant line 10 for directing the oxidant directly to a combustor 6.
5 and an auxiliary oxidant shutoff valve 16 that guides the oxidant to the combustor 6 via the cavitation orifice 17 for the oxidant.
【０００９】図２は、酸化剤タンク圧と燃焼圧との関係
を示す実験結果である。この図において、縦軸は酸化剤
タンク圧、横軸は燃焼圧であり、図中○はキャビテーシ
ョンオリフィスを使用した場合を示し、●はキャビテー
ションオリフィスを使用しない場合を示している。また
図中に示す斜線領域は、振動燃焼が起こる領域と、キャ
ビテーションが起こらない領域である。この図から明ら
かなように、●で示すキャビテーションオリフィスを使
用しない場合には、燃焼圧が高い領域（すなわち、大推
力領域）で安定した燃焼ができ、○で示すキャビテーシ
ョンオリフィスを使用すると、ほぼ同一のタンク圧で比
較的燃焼圧が低い（すなわち、推力の小さい）燃焼を行
うことができる。なお、この実験結果は、燃料タンク圧
と燃焼圧との関係でも同様である。FIG. 2 is an experimental result showing the relationship between the oxidant tank pressure and the combustion pressure. In this figure, the vertical axis represents the oxidant tank pressure, and the horizontal axis represents the combustion pressure. In the figure, ○ indicates the case where the cavitation orifice is used, and ● indicates the case where the cavitation orifice is not used. Further, the shaded areas shown in the figure are an area where oscillating combustion occurs and an area where cavitation does not occur. As is clear from this figure, when the cavitation orifice shown by ● is not used, stable combustion is possible in the high combustion pressure area (that is, the large thrust area), and when the cavitation orifice shown by ○ is used, it is almost the same. Combustion with a relatively low combustion pressure (that is, a small thrust) can be performed with the tank pressure of. The result of this experiment is the same in the relationship between the fuel tank pressure and the combustion pressure.
【００１０】図３は、本発明の推力制御方法を模式的に
示す概念図である。この図において、縦軸は推力であ
り、横軸はタンク圧である。また、図中、上の曲線は主
遮断弁（主燃焼遮断弁１２又は主酸化剤遮断弁１５）を
使用したときのタンク圧と推力の関係を示し、下の曲線
は副遮断弁（副燃焼遮断弁１３又は副酸化剤遮断弁１
６）を使用したときのタンク圧と推力の関係を示してい
る。図２に示したように、キャビテーションオリフィス
の使用の有無により、ほぼ同一のタンク圧力であって
も、主遮断弁と副遮断弁のどちらを用いるかで燃焼圧が
異なり、推力範囲も相違する。また、図中斜線で示す領
域は振動燃焼が起こる領域であり、どちらの場合でもタ
ンク圧力をある値以下に下げると振動燃焼が起こる。し
かし、本発明によれば、大推力時と小推力時とで主遮断
弁と副遮断弁を切替えて使用するので、主遮断弁が適用
できる大推力から副遮断弁が適用できる小推力まで広い
範囲に推力を調節することができる。FIG. 3 is a conceptual diagram schematically showing the thrust control method of the present invention. In this figure, the vertical axis represents thrust and the horizontal axis represents tank pressure. Further, in the figure, the upper curve shows the relationship between the tank pressure and the thrust when the main shutoff valve (main combustion shutoff valve 12 or main oxidant shutoff valve 15) is used, and the lower curve the sub shutoff valve (sub combustion Cutoff valve 13 or auxiliary oxidant cutoff valve 1
It shows the relationship between the tank pressure and thrust when 6) is used. As shown in FIG. 2, depending on whether or not the cavitation orifice is used, even if the tank pressure is almost the same, the combustion pressure and the thrust range differ depending on which of the main shutoff valve and the sub shutoff valve is used. Further, the hatched region in the figure is a region where oscillatory combustion occurs, and in both cases, oscillatory combustion occurs when the tank pressure is lowered below a certain value. However, according to the present invention, since the main shutoff valve and the sub shutoff valve are used by switching between the large thrust and the small thrust, the main shutoff valve can be applied to a large thrust and the sub shutoff valve can be applied to a small thrust. The thrust can be adjusted to the range.
【００１１】上述した二液式ロケットエンジンの推力制
御装置は、以下のように使用される。すなわち、大推力
時には、主燃料遮断弁１２と主酸化剤遮断弁１５を開放
し、副燃料遮断弁１３と副酸化剤遮断弁１６を閉鎖し、
かつ燃料タンク２及び酸化剤タンク４を加圧するガス圧
を調圧器３により制御して、燃焼器６に供給される燃料
及び酸化剤の流量を調節する。この構成は、従来の一般
的な二液式ロケットエンジンの推力制御と同様であり、
大推力時に燃焼器に供給する燃料及び酸化剤の流量を安
定して制御し、燃焼器から噴射する燃焼ガスの流量及び
速度を変化させ、大推力を安定して調節することができ
る。一方、小推力時には、主燃料遮断弁１２と主酸化剤
遮断弁１５を閉鎖し、副燃料遮断弁１３と副酸化剤遮断
弁１６を開放し、かつ、燃料タンク２及び酸化剤タンク
４を加圧するガス圧を制御して、燃焼器に導びかれる燃
料及び酸化剤の流量を調節する。これにより燃料用キャ
ビテーションオリフィス１４及び酸化剤用キャビテーシ
ョンオリフィス１７を介して燃料と酸化剤が燃焼器６に
それぞれ導入される。キャビテーションオリフィスはそ
の縮流部でキャビテーションが発生しており、その部分
の圧力は燃料及び酸化剤の飽和圧力であり一定である。
従って、燃焼器内で振動燃焼が起こり、燃焼器内の燃焼
圧力が大きく変動しても、燃料及び酸化剤の流量は燃焼
器内の燃焼圧力に影響されず上流のガス圧のみで制御す
ることができ、一定の安定した小推力を得ることができ
る。The thrust control device for the above-mentioned two-component rocket engine is used as follows. That is, at the time of large thrust, the main fuel cutoff valve 12 and the main oxidant cutoff valve 15 are opened, the sub fuel cutoff valve 13 and the sub oxidant cutoff valve 16 are closed,
In addition, the gas pressure that pressurizes the fuel tank 2 and the oxidant tank 4 is controlled by the pressure regulator 3 to adjust the flow rates of the fuel and the oxidant supplied to the combustor 6. This configuration is similar to the thrust control of the conventional general two-component rocket engine,
It is possible to stably control the flow rates of the fuel and the oxidizer supplied to the combustor at the time of large thrust, change the flow rate and speed of the combustion gas injected from the combustor, and stably adjust the large thrust. On the other hand, when the thrust is small, the main fuel cutoff valve 12 and the main oxidant cutoff valve 15 are closed, the sub fuel cutoff valve 13 and the sub oxidant cutoff valve 16 are opened, and the fuel tank 2 and the oxidant tank 4 are added. The pressure of the compressed gas is controlled to regulate the flow rate of fuel and oxidant introduced into the combustor. As a result, the fuel and the oxidant are introduced into the combustor 6 through the fuel cavitation orifice 14 and the oxidant cavitation orifice 17, respectively. In the cavitation orifice, cavitation is generated in its contracted portion, and the pressure in that portion is a saturated pressure of the fuel and the oxidizer and is constant.
Therefore, even if oscillatory combustion occurs in the combustor and the combustion pressure in the combustor fluctuates significantly, the flow rates of fuel and oxidant should not be affected by the combustion pressure in the combustor and should be controlled only by the upstream gas pressure. Therefore, a constant and stable small thrust can be obtained.
【００１２】[0012]
【発明の効果】上述したように、本発明によれば、大推
力時には、従来の一般的な二液式ロケットエンジンの推
力制御と同様に、ガス圧の制御により燃焼器に供給する
燃料及び酸化剤の流量を安定して制御し、燃焼器から噴
射する燃焼ガスの流量及び速度を変化させ、大推力を安
定して調節することができ、一方、小推力時には、燃料
用キャビテーションオリフィス及び酸化剤用キャビテー
ションオリフィスを介して燃料と酸化剤を燃焼器に供給
することにより、燃焼器内で振動燃焼が起こり、燃焼器
内の燃焼圧力が大きく変動しても、燃料及び酸化剤の流
量は燃焼器内の燃焼圧力に影響されず上流のガス圧のみ
で制御することができ、一定の安定した小推力を得るこ
とができる。従って、本発明の方法及び装置により、二
液式ロケットエンジンの推力を大推力から小推力まで広
範囲に制御することができる、優れた効果を得ることが
できる。As described above, according to the present invention, at the time of large thrust, the fuel and the oxidant supplied to the combustor are controlled by controlling the gas pressure, as in the thrust control of the conventional general two-liquid rocket engine. It is possible to stably control the flow rate of the agent, change the flow rate and speed of the combustion gas injected from the combustor, and stably adjust the large thrust, while at the time of small thrust, the cavitation orifice for fuel and the oxidizer. By supplying fuel and oxidizer to the combustor through the cavitation orifice for combustion, even if oscillating combustion occurs in the combustor and the combustion pressure in the combustor fluctuates greatly, the flow rates of the fuel and the oxidizer remain constant. It can be controlled only by the upstream gas pressure without being affected by the internal combustion pressure, and a constant and stable small thrust can be obtained. Therefore, with the method and apparatus of the present invention, the thrust of the two-component rocket engine can be controlled in a wide range from large thrust to small thrust, and an excellent effect can be obtained.
【図１】本発明による二液式ロケットエンジンの推力制
御装置の全体構成図である。FIG. 1 is an overall configuration diagram of a thrust control device for a two-component rocket engine according to the present invention.
【図２】酸化剤タンク圧と燃焼圧との関係を示す図であ
る。FIG. 2 is a diagram showing a relationship between an oxidant tank pressure and a combustion pressure.
【図３】本発明の推力制御方法を模式的に示す概念図で
ある。FIG. 3 is a conceptual diagram schematically showing a thrust control method of the present invention.
１ 加圧タンク ２ 燃料タンク ３ 調圧器 ４ 酸化剤タンク ６ 燃焼器 ８ 燃料ライン １０ 酸化剤ライン １２ 主燃料遮断弁 １３ 副燃料遮断弁 １４ 燃料用キャビテーションオリフィス １５ 主酸化剤遮断弁 １６ 副酸化剤遮断弁 １７ 酸化剤用キャビテーションオリフィス 1 Pressurized Tank 2 Fuel Tank 3 Pressure Regulator 4 Oxidizer Tank 6 Combustor 8 Fuel Line 10 Oxidizer Line 12 Main Fuel Cutoff Valve 13 Sub Fuel Cutoff Valve 14 Fuel Cavitation Orifice 15 Main Oxidizer Cutoff Valve 16 Suboxidizer Cutoff Valve 17 Cavitation orifice for oxidizer
Claims (2)
タンクと、燃料を酸化剤で燃焼させる燃焼器と、燃料タ
ンクから燃料を燃焼器に導く燃料ラインと、酸化剤タン
クから酸化剤を燃焼器に導く酸化剤ラインとを備えた二
液式ロケットエンジンの推力制御方法において、 燃料ラインに、燃料を燃焼器に直接導く主燃料遮断弁
と、燃料を燃料用キャビテーションオリフィスを介して
燃焼器に導く副燃料遮断弁とを、互いに並列に設け、 酸化剤ラインに、酸化剤を燃焼器に直接導く主酸化剤遮
断弁と、酸化剤を酸化剤用キャビテーションオリフィス
を介して燃焼器に導く副酸化剤遮断弁とを、互いに並列
に設け、 大推力時に、前記主燃料遮断弁と主酸化剤遮断弁を開放
し、副燃料遮断弁と副酸化剤遮断弁を閉鎖し、 小推力時に、前記主燃料遮断弁と主酸化剤遮断弁を閉鎖
し、副燃料遮断弁と副酸化剤遮断弁を開放し、 かつ、燃料タンク及び酸化剤タンクを加圧するガス圧を
制御して、燃焼器に導びかれる燃料及び酸化剤の流量を
調節する、ことを特徴とする二液式ロケットエンジンの
推力制御方法。1. A fuel tank and an oxidizer tank that are pressurized with gas, a combustor that burns the fuel with the oxidizer, a fuel line that guides the fuel from the fuel tank to the combustor, and an oxidizer from the oxidizer tank. A thrust control method for a two-component rocket engine having an oxidant line leading to a combustor, comprising: a main fuel cutoff valve for directly directing fuel to the combustor to a fuel line; and a combustor for feeding fuel through a cavitation orifice for fuel. A secondary fuel cutoff valve for introducing the oxidant to the combustor is provided in parallel to the oxidant line, and a main oxidant cutoff valve for directly introducing the oxidant to the combustor and a sub fuel cutoff valve for guiding the oxidant to the combustor through the cavitation orifice for the oxidizer. An oxidant shutoff valve is provided in parallel with each other, and at the time of large thrust, the main fuel shutoff valve and the main oxidant shutoff valve are opened, the sub fuel shutoff valve and the suboxidant shutoff valve are closed, and at the time of a small thrust, the Main fuel interception Valve and the main oxidant shutoff valve are closed, the sub fuel shutoff valve and the sub oxidant shutoff valve are opened, and the gas pressure that pressurizes the fuel tank and the oxidant tank is controlled, and the fuel introduced to the combustor is controlled. And a method of controlling thrust of a two-component rocket engine, characterized in that the flow rate of an oxidant is adjusted.
タンクと、燃料を酸化剤で燃焼させる燃焼器と、燃料タ
ンクから燃料を燃焼器に導く燃料ラインと、酸化剤タン
クから酸化剤を燃焼器に導く酸化剤ラインとを備えた二
液式ロケットエンジンの推力制御装置において、 燃料ラインに互いに並列に設けられた、燃料を燃焼器に
直接導く主燃料遮断弁と、燃料を燃料用キャビテーショ
ンオリフィスを介して燃焼器に導く副燃料遮断弁とを備
え、 かつ酸化剤ラインに互いに並列に設けられた、酸化剤を
燃焼器に直接導く主酸化剤遮断弁と、酸化剤を酸化剤用
キャビテーションオリフィスを介して燃焼器に導く副酸
化剤遮断弁とを備え、 前記主燃料遮断弁と主酸化剤遮断弁は、大推力時に開放
し小推力時に閉鎖するように遠隔制御され、 前記副燃料遮断弁と副酸化剤遮断弁は、大推力時に閉鎖
し小推力時に開放するように遠隔制御され、 かつ燃料タンク及び酸化剤タンクを加圧するガス圧は、
燃焼器に導びかれる燃料及び酸化剤の流量を調節するよ
うに制御される、ことを特徴とする二液式ロケットエン
ジンの推力制御装置。2. A fuel tank and an oxidizer tank which are pressurized with gas, a combustor for burning the fuel with the oxidizer, a fuel line for guiding the fuel from the fuel tank to the combustor, and an oxidizer for the oxidizer tank. In a thrust control device for a two-component rocket engine equipped with an oxidant line leading to a combustor, a main fuel cutoff valve for directing fuel directly to a combustor and a fuel cavitation provided in parallel with each other in a fuel line. A main oxidant shutoff valve, which is provided with an auxiliary fuel cutoff valve that leads to the combustor through an orifice, and which is provided in parallel to each other in the oxidant line, and which directly leads the oxidant to the combustor, and cavitation for the oxidant. An auxiliary oxidant shutoff valve that leads to a combustor through an orifice, wherein the main fuel shutoff valve and the main oxidant shutoff valve are remotely controlled so as to open at a large thrust and close at a small thrust. Charge cutoff valve and secondary oxidant shutoff valve is remotely controlled so as to open at the time of a small thrust closed when a large thrust, and the gas pressure for pressurizing the fuel tank and oxidizer tank,
A thrust control device for a two-component rocket engine, wherein the thrust control device is controlled so as to regulate the flow rates of fuel and oxidant introduced to a combustor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP6714393A JPH06280680A (en) | 1993-03-26 | 1993-03-26 | Method and apparatus for controlling thrust of two-liquid system rocket engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP6714393A JPH06280680A (en) | 1993-03-26 | 1993-03-26 | Method and apparatus for controlling thrust of two-liquid system rocket engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH06280680A true JPH06280680A (en) | 1994-10-04 |
Family
ID=13336396
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP6714393A Pending JPH06280680A (en) | 1993-03-26 | 1993-03-26 | Method and apparatus for controlling thrust of two-liquid system rocket engine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH06280680A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2017180461A (en) * | 2016-03-23 | 2017-10-05 | 国立研究開発法人宇宙航空研究開発機構 | Injection system |
| CN112377330A (en) * | 2021-01-18 | 2021-02-19 | 北京星际荣耀空间科技股份有限公司 | Liquid rocket engine thrust adjusting method, device, equipment and storage medium |
-
1993
- 1993-03-26 JP JP6714393A patent/JPH06280680A/en active Pending
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2017180461A (en) * | 2016-03-23 | 2017-10-05 | 国立研究開発法人宇宙航空研究開発機構 | Injection system |
| CN112377330A (en) * | 2021-01-18 | 2021-02-19 | 北京星际荣耀空间科技股份有限公司 | Liquid rocket engine thrust adjusting method, device, equipment and storage medium |
| EP4030047A1 (en) * | 2021-01-18 | 2022-07-20 | Beijing Interstellar Glory Space Technology Co., Ltd. | Thrust adjusting method and apparatus for liquid-propellant rocket engine, device, and storage medium |
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