JPH0972700A - Thrust controller for high-speed airframe - Google Patents
Thrust controller for high-speed airframeInfo
- Publication number
- JPH0972700A JPH0972700A JP22764595A JP22764595A JPH0972700A JP H0972700 A JPH0972700 A JP H0972700A JP 22764595 A JP22764595 A JP 22764595A JP 22764595 A JP22764595 A JP 22764595A JP H0972700 A JPH0972700 A JP H0972700A
- Authority
- JP
- Japan
- Prior art keywords
- stage
- speed
- injection nozzle
- combustion chamber
- control
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Landscapes
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
- Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明は、高速飛翔体の推力
制御装置に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a thrust control device for a high-speed flying object.
【０００２】[0002]
【従来の技術】準中距離以下の弾道ミサイルを大気圏外
で迎撃する高速飛翔体が、米国等で開発されている。図
４は、この迎撃システムを模式的に示すものであり、弾
道ミサイル１の発射が確認されると、これに対応して高
速飛翔体２を発射し、大気圏外で弾道ミサイル１に衝突
させてこれを破壊するようになっている。2. Description of the Related Art A high-speed flying object for intercepting a ballistic missile having a sub-medium range or less outside the atmosphere has been developed in the United States and other countries. FIG. 4 schematically shows the intercepting system. When the launch of the ballistic missile 1 is confirmed, the high-speed flying object 2 is fired in response to this, and the ballistic missile 1 is made to collide with the ballistic missile 1 outside the atmosphere. This is to be destroyed.
【０００３】弾道ミサイル１は、通常多段ロケットであ
り、地上から８０ｋｍ前後の大気圏外では弾道部１ａの
みが曲線３のように慣性飛行し、目標地点まで自由落下
する。これに対して、高速飛翔体２は、例えば３段ロケ
ットであり、（ａ）１段ロケットで大気圏外に達して１
段目を切り離し、（ｂ）２段ロケットで衝突予定地点Ａ
に向けて姿勢制御と速度調整を行った後、２段目を切り
離し、次いで（ｃ）衝突予定付近まで慣性飛行し、
（ｄ）最後に、３段目に設けられた目標捕捉装置で目標
（弾道ミサイル１）を捕らえながら、３段目自体の位置
を制御して弾道ミサイル１に衝突する制御軌跡４を描
く。なお、３段目の飛行自体（ｃ及びｄ）は慣性飛行で
あり、（ｄ）において飛行方向に直交する方向に燃焼ガ
スを噴射して、位置制御するようになっている。[0003] The ballistic missile 1 is usually a multi-stage rocket, and only the ballistic portion 1a inertially flies as shown by a curve 3 outside the atmosphere at about 80 km from the ground and falls freely to a target point. On the other hand, the high-speed flying object 2 is, for example, a three-stage rocket.
Disconnect the stage, and (b) expected collision point A with a two-stage rocket
After performing attitude control and speed adjustment toward, the second stage is separated, and then (c) inertial flight to near the scheduled collision,
(D) Finally, while capturing the target (ballistic missile 1) with the target capturing device provided in the third stage, the position of the third stage itself is controlled to draw the control trajectory 4 that collides with the ballistic missile 1. Note that the third stage flight itself (c and d) is an inertial flight, and the position is controlled by injecting combustion gas in a direction orthogonal to the flight direction in (d).
【０００４】[0004]
【発明が解決しようとする課題】高速飛翔体２の２段目
及び３段目に用いられる燃料として固体推進剤があり、
一旦着火すると、自己燃焼を継続し、その間の燃焼速度
の調整は困難である。そのため、従来の２段目の燃焼
（ｂ）では、燃料の内面形状を工夫してできるだけ一定
の燃焼速度にすると共に、その燃焼の継続中に、姿勢制
御と速度制御を完了する必要がある。しかし、姿勢制御
は、ノズルの噴射方向の制御等により比較的容易である
が、速度制御は、短時間に所定の高速に達するように
内圧を高めると、空気抵抗が低下した大気圏外では燃料
の残量により更に加速されるおそれがあり、逆に所定
の速度で燃料がほぼ尽きるように徐々に加速すると、所
定の速度に達するのに時間がかかり、弾道ミサイル１を
確実に迎撃できなくなる問題点があった。There is a solid propellant as a fuel used in the second and third stages of the high-speed projectile 2.
Once ignited, self-combustion continues and it is difficult to adjust the burning rate during that period. Therefore, in the conventional second-stage combustion (b), it is necessary to devise the shape of the inner surface of the fuel to make the combustion speed as constant as possible, and to complete the attitude control and the speed control while continuing the combustion. However, the attitude control is relatively easy by controlling the injection direction of the nozzles, etc., but the speed control is such that if the internal pressure is increased to reach a predetermined high speed in a short time, the fuel resistance will be increased outside the atmosphere where the air resistance has decreased. There is a risk of further acceleration depending on the remaining amount. Conversely, if the fuel is gradually accelerated at a predetermined speed so that the fuel is almost exhausted, it will take time to reach the predetermined speed, and it will not be possible to reliably intercept the ballistic missile 1. was there.
【０００５】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、短時
間に所定の高速に達することができ、かつ燃料の残量に
よる加速がなく、所定の速度をそのまま保持することが
できる高速飛翔体の推力制御装置を提供することにあ
る。The present invention was devised to solve such problems. That is, an object of the present invention is to provide a thrust control device for a high-speed flying vehicle, which can reach a predetermined high speed in a short time, does not accelerate due to the remaining amount of fuel, and can maintain the predetermined speed as it is. Especially.
【０００６】[0006]
【課題を解決するための手段】本発明によれば、第１固
体推進剤を内蔵する第１燃焼室と推進用噴射ノズルを有
する２段目と、第２固体推進剤を内蔵する第２燃焼室と
姿勢制御用噴射ノズルを有する３段目と、を備えた高速
飛翔体の推力制御装置において、２段目の第１燃焼室か
ら３段目の姿勢制御用噴射ノズルまで燃焼ガスを導く放
出ガス流路と、該放出ガス流路を通るガス流量を調節す
る流量調節弁と、第１燃焼室内の圧力を検出して流量調
節弁を制御する制御装置と、を備えたことを特徴とする
高速飛翔体の推力制御装置が提供される。According to the present invention, a first combustion chamber containing a first solid propellant and a second stage having a propelling injection nozzle, and a second combustion containing a second solid propellant. In a thrust control device for a high-speed projectile, which includes a chamber and a third stage having an attitude control injection nozzle, a discharge for guiding combustion gas from the first combustion chamber in the second stage to the injection nozzle for attitude control in the third stage A gas flow path, a flow rate control valve for controlling a gas flow rate through the discharge gas flow channel, and a control device for detecting the pressure in the first combustion chamber to control the flow rate control valve. A thrust control device for a high-speed flying object is provided.
【０００７】固体推進剤の燃焼速度ｒは、ｒ＝ａ・ｐn
（式１）の実験式で表すことができる。ここで、ａは定
数、ｐは燃焼圧、ｎは推進剤による定数である。従っ
て、上記本発明を構成する制御装置により第１燃焼室内
の圧力を検出して流量調節弁を制御することにより、目
標の速度に達するまでは内圧を高めて加速時間を短縮
し、次いで所定の速度に達した後は、一定速度を保持す
るように内圧を制御することができる。The burning velocity r of the solid propellant is r = a · p n
It can be expressed by the empirical formula of (Formula 1). Here, a is a constant, p is a combustion pressure, and n is a constant depending on the propellant. Therefore, by detecting the pressure in the first combustion chamber and controlling the flow rate control valve by the control device constituting the present invention, the internal pressure is increased to shorten the acceleration time until the target speed is reached, and then the predetermined time is reached. After reaching the speed, the internal pressure can be controlled to maintain a constant speed.
【０００８】また、流量調節弁を通して姿勢制御用噴射
ノズルに導かれた燃焼ガスは、複数の噴射ノズルから反
対方向に同量噴射させれば、姿勢に影響なく外部に放出
することができ、或いは複数の噴射ノズルの噴出量を調
整すれば、２段目の燃焼時にも姿勢制御に役立てること
ができ、２段目の方向制御機構等を簡略化することがで
きる。The combustion gas guided to the attitude control injection nozzle through the flow rate control valve can be discharged to the outside without affecting the attitude by injecting the same amount in the opposite direction from a plurality of injection nozzles, or If the ejection amounts of the plurality of injection nozzles are adjusted, the attitude control can be utilized even during the second stage combustion, and the second direction control mechanism and the like can be simplified.
【０００９】更に、本発明の好ましい実施形態によれ
ば、前記第１及び第２固体推進剤は、軸線に沿った貫通
孔をそれぞれ有する。この構成により、第１固体推進剤
の軸方向反対側に推進用噴射ノズルと放出ガス流路を配
置することができ、かつ第２固体推進剤の軸方向反対側
に姿勢制御用噴射ノズルを別々に配置することができ
る。Further, according to a preferred embodiment of the present invention, the first and second solid propellants each have a through hole extending along an axis. With this configuration, the propulsion injection nozzle and the discharge gas passage can be arranged on the axially opposite side of the first solid propellant, and the attitude control injection nozzle is separately provided on the axially opposite side of the second solid propellant. Can be placed at.
【００１０】[0010]
【発明の実施の形態】以下、本発明の好ましい実施形態
を図面を参照して説明する。なお、各図において共通す
る部分には同一の符号を付して使用する。図１は、本発
明による推力制御装置を備えた高速飛翔体の構成図であ
る。この図に示すように、この高速飛翔体は、第１固体
推進剤５ａを内蔵する第１燃焼室５と推進用噴射ノズル
６を有する２段目部分７と、第２固体推進剤８ａを内蔵
する第２燃焼室８と姿勢制御用噴射ノズル９を有する３
段目部分１０と、を備えている。更に、この高速飛翔体
は、１段目部分１１を有し、この１段目部分１１の推進
ロケットにより、図４に例示したように大気圏外まで達
することができるようになっている。DESCRIPTION OF THE PREFERRED EMBODIMENTS Preferred embodiments of the present invention will be described below with reference to the drawings. In the drawings, common parts are denoted by the same reference numerals. FIG. 1 is a configuration diagram of a high-speed flying object including a thrust control device according to the present invention. As shown in this figure, this high-speed projectile contains a first combustion chamber 5 containing a first solid propellant 5a, a second stage portion 7 having a propulsion injection nozzle 6, and a second solid propellant 8a. 3 having the second combustion chamber 8 and the attitude control injection nozzle 9
And a step portion 10. Further, this high-speed flying object has a first-stage portion 11, and the propulsion rocket of the first-stage portion 11 can reach outside the atmosphere as illustrated in FIG.
【００１１】固体推進剤５ａ，８ａは、固体ロケット用
の固体推進剤であり、分子内に燃焼を維持するのに十分
な酸素を含む物質からなる均質系推薬であっても、或い
は酸化剤を別に混合した混合系推薬であってもよい。ま
た、固体推進剤５ａ，８ａには、図に示すように、軸線
に沿った貫通孔がそれぞれ設けられている。推進用噴射
ノズル６は、２段目部分７の下部に設けられ、第１固体
推進剤５ａの燃焼ガス（反応ガス）を下方に噴出し推力
を発生するようになっている。この推進用噴射ノズル６
は、噴出方向を調整して姿勢制御ができるようになって
いるのがよい。The solid propellants 5a and 8a are solid propellants for solid rockets and may be homogeneous propellants made of a substance containing sufficient oxygen in the molecule to maintain combustion, or oxidizers. It may be a mixed propellant in which Further, as shown in the figure, the solid propellants 5a and 8a are respectively provided with through holes along the axis. The propulsion injection nozzle 6 is provided in the lower part of the second stage portion 7, and is configured to eject the combustion gas (reaction gas) of the first solid propellant 5a downward to generate thrust. This propulsion injection nozzle 6
It is good that the posture can be controlled by adjusting the ejection direction.
【００１２】姿勢制御用噴射ノズル９は、第２燃焼室８
の上下に別れた複数の大噴射ノズル９ａと複数の小噴射
ノズル９ｂからなる。各噴射ノズル９ａ，９ｂは、それ
ぞれ軸線に対して直交する方向に燃焼ガスを噴射する。
また、各噴射ノズル９ａ，９ｂには流量調節弁１４ａと
開閉弁付流量調節弁１４ｂがそれぞれ取り付けられ、こ
の流量調節弁１４ａ，１４ｂにより各ノズルからの噴射
量を調節し、３段目部分１０の姿勢制御を行うようにな
っている。かかる構成により、姿勢制御用噴射ノズル９
により推力を発生させることなく３段目部分１０の姿勢
制御を行うことができる。The attitude control injection nozzle 9 is provided in the second combustion chamber 8
It is composed of a plurality of large jet nozzles 9a and a plurality of small jet nozzles 9b which are separated from each other. Each of the injection nozzles 9a and 9b injects combustion gas in a direction orthogonal to the axis.
A flow rate adjusting valve 14a and a flow rate adjusting valve 14b with an on-off valve are attached to each of the injection nozzles 9a and 9b. The flow rate adjusting valves 14a and 14b are used to adjust the injection amount from each nozzle to adjust the third stage portion 10 Attitude control. With this configuration, the attitude control injection nozzle 9
Thus, the attitude control of the third stage portion 10 can be performed without generating thrust.
【００１３】更に、第２固体推進剤８ａの中央部に設け
られた貫通孔と、噴射ノズル９ａ，９ｂとを連通するガ
ス通路には、開閉弁１２が設けられており、この開閉弁
１２により、第２燃焼室８で発生した燃焼ガスの噴射ノ
ズル９ａ，９ｂへの供給の開閉できるようになってい
る。上述した構成により、第１固体推進剤５ａの軸方向
反対側に推進用噴射ノズル６と放出ガス流路２２を配置
することができ、かつ第２固体推進剤８ａの軸方向に噴
射ノズル９ａ，９ｂを別々に配置することができる。Further, an on-off valve 12 is provided in the gas passage that connects the through hole provided in the central portion of the second solid propellant 8a and the injection nozzles 9a and 9b. The supply of the combustion gas generated in the second combustion chamber 8 to the injection nozzles 9a and 9b can be opened and closed. With the configuration described above, the propulsion injection nozzle 6 and the discharge gas flow path 22 can be arranged on the axially opposite side of the first solid propellant 5a, and the injection nozzle 9a, 9a, 9b can be arranged separately.
【００１４】本発明の推力制御装置２０は、２段目の第
１燃焼室５から３段目の姿勢制御用噴射ノズル９まで燃
焼ガスを導く放出ガス流路２２と、放出ガス流路２２を
通るガス流量を調節する流量調節弁２４と、第１燃焼室
５内の圧力を検出して流量調節弁２４を制御する制御装
置２６と、を備えている。The thrust control device 20 of the present invention includes a discharge gas passage 22 for guiding combustion gas from the first combustion chamber 5 in the second stage to the injection nozzle 9 for posture control in the third stage, and a discharge gas passage 22. A flow rate control valve 24 that controls the flow rate of the gas passing through and a control device 26 that detects the pressure in the first combustion chamber 5 and controls the flow rate control valve 24 are provided.
【００１５】図２は、上述した高速飛翔体の推力制御を
示す模式図（図４のｂの状態に相当する）であり、図３
は、図４のｄにおける位置制御状態を示す模式図であ
る。上述したように、固体推進剤の燃焼速度ｒは、ｒ＝
ａ・ｐn （式１）の実験式で表すことができる（ａは定
数、ｐは燃焼圧、ｎは推進剤による定数）ので、図２に
示すように、制御装置２６により第１燃焼室５内の圧力
を検出して流量調節弁２４を制御することにより、目標
の速度に達するまでは内圧を高めて加速時間を短縮し、
次いで所定の速度に達した後は、一定速度を保持するよ
うに内圧を制御することができる。また、流量調節弁２
４を通して姿勢制御用噴射ノズル９ｂに導かれた燃焼ガ
スは、複数の噴射ノズルから反対方向に同量噴射させれ
ば、姿勢に影響なく外部に放出することができ、或いは
複数の噴射ノズルの噴出量を調整すれば、２段目の燃焼
時にも姿勢制御に役立てることができ、２段目の方向制
御機構等を簡略化することができる。なお、図２に示す
状態において、開閉弁付流量調節弁１４ｂは閉じてお
り、第２固体推進剤８ａは未着火であることは勿論であ
る。FIG. 2 is a schematic diagram (corresponding to the state of b in FIG. 4) showing the thrust control of the above-mentioned high-speed flying body, and FIG.
FIG. 5 is a schematic diagram showing a position control state in d of FIG. As described above, the burning velocity r of the solid propellant is r =
Since it can be expressed by an empirical formula of a · pn (Equation 1) (a is a constant, p is a combustion pressure, and n is a constant depending on a propellant), the control device 26 controls the first combustion chamber as shown in FIG. By detecting the pressure in 5 and controlling the flow control valve 24, the internal pressure is increased and the acceleration time is shortened until the target speed is reached.
Then, after reaching a predetermined speed, the internal pressure can be controlled so as to maintain a constant speed. Also, the flow control valve 2
The combustion gas guided to the attitude control injection nozzle 9b through 4 can be discharged to the outside without affecting the attitude by injecting the same amount in the opposite direction from the plurality of injection nozzles, or the ejection of the plurality of injection nozzles. If the amount is adjusted, it can be useful for posture control even during the second stage combustion, and the second stage direction control mechanism and the like can be simplified. In addition, in the state shown in FIG. 2, it goes without saying that the flow rate control valve with an on-off valve 14b is closed and the second solid propellant 8a has not been ignited.
【００１６】次いで、図３（図４のｄの状態）におい
て、流量調節弁２４は閉じ、開閉弁付流量調節弁１４ｂ
が開き、第２固体推進剤８ａの燃焼により、噴射ノズル
９ａ，９ｂを通して飛行方向に直交する方向に燃焼ガス
を噴射して、３段目部分の位置制御を行う。なお、この
状態では２段目は切り離されており、噴射ノズル９ａ，
９ｂでは推力を発生させず、慣性飛行を続ける。また、
噴射ノズル９ａ，９ｂの一部を斜め後方に向けて取り付
け、図３において推力を発生させて加速できるようにし
てもよい。Next, in FIG. 3 (state of FIG. 4d), the flow rate control valve 24 is closed and the flow rate control valve 14b with an opening / closing valve is closed.
Opens, and the combustion gas of the second solid propellant 8a is injected to inject the combustion gas through the injection nozzles 9a and 9b in a direction orthogonal to the flight direction to control the position of the third stage. In this state, the second stage is separated, and the injection nozzle 9a,
At 9b, thrust is not generated and inertial flight continues. Also,
A part of the injection nozzles 9a and 9b may be attached obliquely rearward so as to generate thrust in FIG. 3 so as to be accelerated.
【００１７】なお、本発明は上述した実施形態に限定さ
れず、本発明の要旨を逸脱しない範囲で種々に変更でき
ることは勿論である。It should be noted that the present invention is not limited to the above-described embodiment, but can be variously modified without departing from the gist of the present invention.
【００１８】[0018]
【発明の効果】上述したように、本発明の高速飛翔体の
推力制御装置は、短時間に所定の高速に達することがで
き、かつ燃料の残量による加速がなく、所定の速度をそ
のまま保持することができる、等の優れた効果を有す
る。As described above, the thrust control device for a high-speed flying object of the present invention can reach a predetermined high speed in a short time and does not accelerate due to the remaining amount of fuel and keeps the predetermined speed as it is. It has an excellent effect such as being able to.
【図１】本発明による推力制御装置を備えた高速飛翔体
の構成図である。FIG. 1 is a configuration diagram of a high-speed flying object including a thrust control device according to the present invention.
【図２】高速飛翔体の推力制御を示す模式図である。FIG. 2 is a schematic diagram showing thrust control of a high-speed flying object.
【図３】図４のｄにおける位置制御状態を示す模式図で
ある。FIG. 3 is a schematic diagram showing a position control state in d of FIG.
【図４】従来の迎撃システムの模式図である。FIG. 4 is a schematic view of a conventional interception system.
１ 弾道ミサイル １ａ 弾道部 ２ 高速飛翔体 ３ 慣性飛行曲線 ４ 制御軌跡 ５ 第１燃焼室 ５ａ 第１固体推進剤 ６ 推進用噴射ノズル ７ ２段目部分 ８ 第２燃焼室 ８ａ 第２固体推進剤 ９ 姿勢制御用噴射ノズル ９ａ 大噴射ノズル ９ｂ 小噴射ノズル １０ ３段目部分 １１ １段目部分 １２ 開閉弁 １４ａ 流量調節弁 １４ｂ 開閉弁付流量調節弁 ２０ 推力制御装置 ２２ 放出ガス流路 ２４ 流量調節弁 ２６ 制御装置 1 Ballistic Missile 1a Ballistic Part 2 High-speed Flying Body 3 Inertial Flight Curve 4 Control Trajectory 5 1st Combustion Chamber 5a 1st Solid Propellant 6 Propulsion Injection Nozzle 7 2nd Stage 8 2nd Combustion Chamber 8a 2nd Solid Propellant 9 Posture control injection nozzle 9a Large injection nozzle 9b Small injection nozzle 10 Third stage part 11 First stage part 12 Open / close valve 14a Flow control valve 14b Flow control valve with open / close valve 20 Thrust control device 22 Emission gas flow path 24 Flow control valve 26 Control device
Claims (2)
推進用噴射ノズルを有する２段目と、第２固体推進剤を
内蔵する第２燃焼室と姿勢制御用噴射ノズルを有する３
段目と、を備えた高速飛翔体の推力制御装置において、 ２段目の第１燃焼室から３段目の姿勢制御用噴射ノズル
まで燃焼ガスを導く放出ガス流路と、該放出ガス流路を
通るガス流量を調節する流量調節弁と、第１燃焼室内の
圧力を検出して流量調節弁を制御する制御装置と、を備
えたことを特徴とする高速飛翔体の推力制御装置。1. A second stage having a first combustion chamber containing a first solid propellant and a propulsion injection nozzle, and a second stage having a second combustion chamber containing a second solid propellant and an attitude control injection nozzle.
In a thrust control device for a high-speed projectile including a stage, a discharge gas flow path for guiding combustion gas from a first combustion chamber in the second stage to an injection nozzle for attitude control in the third stage, and the discharge gas flow channel A thrust control device for a high-speed projectile, comprising: a flow rate control valve for controlling the flow rate of gas passing through the first combustion chamber; and a control device for detecting the pressure in the first combustion chamber to control the flow rate control valve.
沿った貫通孔をそれぞれ有する、ことを特徴とする請求
項１に記載の高速飛翔体の推力制御装置。2. The thrust control device for a high-speed projectile according to claim 1, wherein the first and second solid propellants each have a through hole along an axis.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22764595A JPH0972700A (en) | 1995-09-05 | 1995-09-05 | Thrust controller for high-speed airframe |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22764595A JPH0972700A (en) | 1995-09-05 | 1995-09-05 | Thrust controller for high-speed airframe |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH0972700A true JPH0972700A (en) | 1997-03-18 |
Family
ID=16864127
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP22764595A Pending JPH0972700A (en) | 1995-09-05 | 1995-09-05 | Thrust controller for high-speed airframe |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH0972700A (en) |
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|---|---|---|---|---|
| JP2009257154A (en) * | 2008-04-15 | 2009-11-05 | Ihi Aerospace Co Ltd | Side thruster of flying object |
| JP2010503794A (en) * | 2006-09-13 | 2010-02-04 | アエロジェット ジェネラル コーポレイション | Pintle control propulsion system with external dynamic seal |
| CN101892925A (en) * | 2010-06-23 | 2010-11-24 | 中北大学 | Aircraft range-extending technique |
| JP2011236903A (en) * | 2010-05-11 | 2011-11-24 | Alliant Techsyst Inc | Rockets, methods of rocket control and methods of rocket evaluation utilizing pressure compensation |
| JP2013007328A (en) * | 2011-06-24 | 2013-01-10 | Ihi Aerospace Co Ltd | Pulse rocket motor and missile |
| JP2014517243A (en) * | 2011-05-19 | 2014-07-17 | エラクレス | System for thrust steering and attitude control during flight of vehicle, and aircraft equipped with the system |
| WO2017072457A1 (en) * | 2015-10-28 | 2017-05-04 | Airbus Safran Launchers Sas | More compact direct thrust flight control and attitude control system, and craft comprising such a system |
| CN113028453A (en) * | 2021-04-09 | 2021-06-25 | 西北工业大学 | Rotary detonation combustion chamber with adjustable combustion chamber width |
-
1995
- 1995-09-05 JP JP22764595A patent/JPH0972700A/en active Pending
Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2010503794A (en) * | 2006-09-13 | 2010-02-04 | アエロジェット ジェネラル コーポレイション | Pintle control propulsion system with external dynamic seal |
| JP2009257154A (en) * | 2008-04-15 | 2009-11-05 | Ihi Aerospace Co Ltd | Side thruster of flying object |
| JP2011236903A (en) * | 2010-05-11 | 2011-11-24 | Alliant Techsyst Inc | Rockets, methods of rocket control and methods of rocket evaluation utilizing pressure compensation |
| CN101892925A (en) * | 2010-06-23 | 2010-11-24 | 中北大学 | Aircraft range-extending technique |
| JP2014517243A (en) * | 2011-05-19 | 2014-07-17 | エラクレス | System for thrust steering and attitude control during flight of vehicle, and aircraft equipped with the system |
| JP2013007328A (en) * | 2011-06-24 | 2013-01-10 | Ihi Aerospace Co Ltd | Pulse rocket motor and missile |
| WO2017072457A1 (en) * | 2015-10-28 | 2017-05-04 | Airbus Safran Launchers Sas | More compact direct thrust flight control and attitude control system, and craft comprising such a system |
| FR3043067A1 (en) * | 2015-10-28 | 2017-05-05 | Herakles | POWER STEERING AND ATTITUDE CONTROL SYSTEM WITH INCREASED COMPACITY AND FINALLY HAVING SUCH A SYSTEM |
| KR20180100307A (en) * | 2015-10-28 | 2018-09-10 | 아리안그룹 에스아에스 | A small lateral force steering and attitude control system, and a vehicle including the system |
| US10895443B2 (en) | 2015-10-28 | 2021-01-19 | Arianegroup Sas | More compact side force steering and attitude control system, and a vehicle including such a system |
| CN113028453A (en) * | 2021-04-09 | 2021-06-25 | 西北工业大学 | Rotary detonation combustion chamber with adjustable combustion chamber width |
| CN113028453B (en) * | 2021-04-09 | 2022-07-01 | 西北工业大学 | Rotary detonation combustion chamber with adjustable combustion chamber width |
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