JPH0319907B2 - - Google Patents
Info
- Publication number
- JPH0319907B2 JPH0319907B2 JP58001885A JP188583A JPH0319907B2 JP H0319907 B2 JPH0319907 B2 JP H0319907B2 JP 58001885 A JP58001885 A JP 58001885A JP 188583 A JP188583 A JP 188583A JP H0319907 B2 JPH0319907 B2 JP H0319907B2
- Authority
- JP
- Japan
- Prior art keywords
- combustion
- rocket
- chamber
- thrust
- rocket chamber
- 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.)
- Expired - Lifetime
Links
- 238000002485 combustion reaction Methods 0.000 claims description 56
- 239000000463 material Substances 0.000 claims description 26
- 239000007779 soft material Substances 0.000 claims description 2
- 230000002250 progressing effect Effects 0.000 description 3
- 230000000903 blocking effect Effects 0.000 description 2
- 235000015842 Hesperis Nutrition 0.000 description 1
- 235000012633 Iberis amara Nutrition 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000003380 propellant Substances 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/26—Burning control
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
Description
【発明の詳細な説明】
本発明は高々度飛翔体として好適なロケツトエ
ンジンに関するものである。DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a rocket engine suitable for use as a high-altitude flying vehicle.
周知のごとく、高々度に飛翔体を効果的に到達
させる様打上げるには、そのロケツトエンジンの
推力と推力の作用時間（通常、作動時間という）
の最適の組合せがある。（特許出願公告昭41−762
号、特許出願公告昭41−6762号）。 As is well known, in order to effectively launch a projectile to a high altitude, the thrust of the rocket engine and the operating time of the thrust (usually referred to as operating time) are essential.
There is an optimal combination of (Patent application announcement 1976-762
No., Patent Application Publication No. 1973-6762).
併し乍ら、従来のロケツトは最適な推力×時間
の組合せが得られない内面燃焼形式が殆んどであ
つた。また、端面燃焼形式では推力が不足で効率
が悪い。一部上記特許願による高燃焼速度推進剤
等を軸上に埋め込んだ特殊端面燃焼形式もあるが
広範囲に使用されてはいない。この点に関して最
適な推力×時間のロケツトエンジンを得るため
種々の試みがなされて来たが、いずれも十分なも
のではなかつた。 However, most conventional rockets have been of the internal combustion type, which does not provide the optimum thrust x time combination. Additionally, the end-burning type lacks thrust and is inefficient. There is also a special end-combustion type in which a high-burning-rate propellant, etc. is embedded on the shaft, as disclosed in some of the above-mentioned patent applications, but it is not widely used. In this regard, various attempts have been made to obtain a rocket engine with optimal thrust x time, but none of them have been satisfactory.
本発明は、現在容易に得られる燃焼速度のグレ
インのみを用いて最適な推力×時間のロケツトエ
ンジンを得ることを目的とし、具体的にはロケツ
トチヤンバーに装填されたグレインの円周方向に
部分的に配設され、その一部が外周へ閉じている
様な燃焼制限材を、ロケツトチヤンバーの軸方向
に沿つて配設する構造によつてその目的を達する
ものである。そして、さらには上記燃焼制限材を
ロケツトチヤンバーの軸方向に沿つて部分的に配
設することにより、さらに良好な推力を得ようと
するものである。 The purpose of the present invention is to obtain a rocket engine with an optimal thrust x time using only grains with a burning rate that can be easily obtained at present. This objective is achieved by a structure in which a combustion restricting material is disposed along the axial direction of the rocket chamber, and a part of the combustion restricting material is closed to the outer periphery. Further, by arranging the combustion restricting material partially along the axial direction of the rocket chamber, it is attempted to obtain even better thrust.
即ち本発明は、ロケツトチヤンバーに内面燃焼
形式のグレインが装填され、該グレインはロケツ
トチヤンバーの中心軸方向に延びる複数個の板状
燃焼制限材で区画されており、これらの燃焼制限
材はロケツトの中心軸に対し垂直な横断面に於い
てロケツトチヤンバーの内周から離隔した円周に
沿つて相互に間隔をおいて配設された円周部分
と、これらの円周部分を夫々ロケツトチヤンバー
の内周へ閉じる連結部分とからなることを特徴と
するロケツトエンジンに係るものである。 That is, in the present invention, a rocket chamber is loaded with internal combustion type grains, and the grains are partitioned by a plurality of plate-shaped combustion restriction materials extending in the direction of the central axis of the rocket chamber, and these combustion restriction materials are circumferential portions spaced from each other along a circumference spaced apart from the inner circumference of the rocket chamber in a cross section perpendicular to the central axis of the rocket; This invention relates to a rocket engine characterized by comprising a connecting portion that closes to the inner periphery of a chamber.
本発明によれば、かかる構造を有する燃焼制限
材を用いることにより内面燃焼形式のグレインに
おける燃焼遂行距離（以下ウエブと言う）を実質
上増加させ得る。 According to the present invention, by using a combustion restricting material having such a structure, the combustion distance (hereinafter referred to as "web") in internal combustion type grains can be substantially increased.
即ち燃焼の進行が初期段階ではロケツトエンジ
ン横断面の略中心部からあらゆる放射方向へ進む
が、燃焼制限材の各円周部分の内面に達するとこ
れら円周部分の間隙にある成分のみが燃焼を続け
て燃焼進行面がロケツトチヤンバーに達し、その
後燃焼制限材の円周部分の外側に沿つて円周方向
に燃焼を続け、これらの円周部分をロケツトチヤ
ンバーの内周へ閉じる連結部分に於いて燃焼の進
行が終わる。本発明に於いて燃焼制限材の円周部
分を閉じるとは、円周部分の一方の端或いは中間
点をロケツトチヤンバーにくつつけて燃焼を遮断
することを意味する。この連結部分は燃焼の進行
を遮断する作用と共に燃焼末期において燃焼制限
材をロケツトチヤンバー内で支持する作用をも有
する。 In other words, in the initial stage, combustion progresses in all radial directions from approximately the center of the cross-section of the rocket engine, but when it reaches the inner surface of each circumferential portion of the combustion restriction material, only the components in the gaps between these circumferential portions are combusted. The combustion progressing surface then reaches the rocket chamber, and then continues burning in the circumferential direction along the outside of the circumferential portion of the combustion restriction material until it reaches the connecting portion that closes these circumferential portions to the inner circumference of the rocket chamber. The combustion process ends at this point. In the present invention, closing the circumferential portion of the combustion restricting material means attaching one end or midpoint of the circumferential portion to the rocket chamber to shut off combustion. This connecting portion has the function of blocking the progress of combustion as well as the function of supporting the combustion restricting material within the rocket chamber at the final stage of combustion.
以下、本発明の一実施例を図面に基づいて説明
する。 Hereinafter, one embodiment of the present invention will be described based on the drawings.
本発明の一実施例の断面形状を示すと第１図か
ら第３図示す様になる。第１図及びび第２図では
円周に沿つて配設された三つの燃焼制限材３が
夫々一端に於いてロケツトチヤンバー４へ閉じた
形状を有し、第３図は二つの燃焼制限材３が夫々
中央に於いてロケツトチヤンバーへ閉じた形状を
有する。通常、内面燃焼形式によるウエブはその
グレインの直径の25％〜30％であるのに対し、本
発明においてはグレイン１の円周方向に部分的に
配設された燃焼制限材３によつて実質上通常のグ
レイン１の直径の２乃至３倍のウエブが得られる
ようになる。このうち第３図に示した実施例につ
いて燃焼進行ウエブを実質的に増加させる効果を
説明すると、第４図に示す如くになる。この例で
は約2.5倍に増加している。即ち第４図に於いて
１ａはグレイン燃焼初期面を示し、Ａは燃焼進行
面、３は燃焼制限材、４はロケツトチヤンバーを
示す。燃焼制限材３は円周に沿つて配設された円
周部分がとの中央に於いて連結部分３′によりロ
ケツトチヤンバー４の内周にくつつけられて閉じ
た形状を有する。 The cross-sectional shape of one embodiment of the present invention is as shown in FIGS. 1 to 3. 1 and 2, three combustion restriction members 3 disposed along the circumference each have a shape closed to the rocket chamber 4 at one end, and FIG. 3 shows two combustion restriction members 3. The members 3 each have a closed shape in the center into the rocket chamber. Normally, the internal combustion type web is 25% to 30% of the grain diameter, but in the present invention, the combustion restriction material 3 partially disposed in the circumferential direction of the grain 1 substantially A web with a diameter two to three times the diameter of the normal grain 1 can be obtained. Among these, the effect of substantially increasing the combustion progressing web in the embodiment shown in FIG. 3 will be explained as shown in FIG. 4. In this example, it has increased by about 2.5 times. That is, in FIG. 4, 1a indicates the initial grain combustion surface, A indicates the combustion progressing surface, 3 indicates the combustion restriction material, and 4 indicates the rocket chamber. The combustion restricting material 3 has a closed shape in which a circumferential portion disposed along the circumference is attached to the inner periphery of the rocket chamber 4 by a connecting portion 3' at the center thereof.
第４図に於いては燃焼は初期面１ａからＡで指
示する如く進行する。図示の如く初期段階ではあ
らゆる放射方向に進むが、そのうち燃焼制限材３
に妨げられない上下の２方向の成分の燃焼がロケ
ツトチヤンバーに達した後、燃焼制限材の外側に
於いてチヤンバーの円周方向に進行し、燃焼制限
材３が外周に閉じる連結部分３′で最終状態に到
達する。従つて、ウエブ（燃焼進行距離）は第４
図に於いてC1＋C2であり、燃焼制限材３のない
場合の2.5倍である。この例による燃焼パターン
を示すと第５図に示す様になる。即ち、燃焼初期
の段階に於ては比較的大きな推力Ｆが得られ、一
定時間が経過した後は比較的小さなほぼ一定の安
定した推力が継続して得られる。この様な推力×
時間の組合せについては、燃焼制限材３の位置・
形状を変化させることで種々の設計変更が可能で
ある。本発明を用いてロケツトエンジンを設計す
れば、初期の推力を充分大きく出来、又その後も
所望の推力が得られる為、従来行なわれていた様
にランチヤー離脱時の速度を補う為にブースタグ
レインを併用することなどは必要でない。要求に
よつて燃焼制限材３をロケツトチヤンバー軸方向
に沿つた全長に亘つて施さず、例おば第６図に示
すように部分的にのみ配設すれば、初期の推力を
さらに増大させることが出来る。尚、第６図中５
はノズルを示す。即ち第６図に於いてはロケツト
エンジン２のロケツトチヤンバー４の軸方向のノ
ズル５に近い部分を除いた内方部分にのみ燃焼制
限材３が配設されている。さらに、グレイン１内
に配設される燃焼制限材はその材質を軟質なもの
とするか、或いは二層式又は多層式構造とするこ
とにより温度環境或いは取り扱い等によつて生ず
る応力を緩和し、グレイン１の破壊又は内在応力
ひずみを防ぐことが出来る。その実施例を第７図
に示す。 In FIG. 4, combustion proceeds as indicated by A from the initial surface 1a. As shown in the figure, in the initial stage, the flame propagates in all radial directions, but eventually the flame-restricting material 3
After the combustion of the components in the upper and lower directions, which are not hindered by the reaches the final state. Therefore, the web (burning progress distance) is the fourth
In the figure, it is C 1 +C 2 , which is 2.5 times the case without the combustion restriction material 3. The combustion pattern according to this example is shown in FIG. That is, in the initial stage of combustion, a relatively large thrust F is obtained, and after a certain period of time, a relatively small, almost constant, stable thrust is continuously obtained. This kind of thrust ×
Regarding the combination of time, the position of combustion restriction material 3 and
Various design changes are possible by changing the shape. If a rocket engine is designed using the present invention, the initial thrust can be sufficiently large and the desired thrust can be obtained thereafter, so booster grains can be used to compensate for the speed when the launcher leaves the rocket, as was conventionally done. It is not necessary to use them together. If the combustion restriction material 3 is not applied along the entire length of the rocket chamber along the axial direction, but only partially as shown in FIG. 6, the initial thrust can be further increased. I can do it. In addition, 5 in Figure 6
indicates a nozzle. That is, in FIG. 6, the combustion restricting material 3 is disposed only in the inner part of the rocket chamber 4 of the rocket engine 2, excluding the part near the nozzle 5 in the axial direction. Furthermore, the combustion restricting material disposed within the grain 1 is made of a soft material or has a two-layer or multi-layer structure to alleviate stress caused by the temperature environment or handling, etc. Breaking of the grain 1 or inherent stress strain can be prevented. An example thereof is shown in FIG.
本発明は以上の如く構成されている為、内面燃
焼形式を用いた場合と同様の初期の充分な推力が
得られると共に、その後はほぼ一定した推力を継
続して得ることが出来、従つてこれによつて推力
×時間の関係が良好な燃焼パターンの設計が容易
となる。又、燃焼制限材を軟質或いは二層式・又
は多層式構造とすることにより、グレイン１の破
壊やひずみ等をもたらすこともなくなる。 Since the present invention is configured as described above, it is possible to obtain a sufficient initial thrust similar to that when using the internal combustion type, and after that, it is possible to continuously obtain a substantially constant thrust. This makes it easy to design a combustion pattern with a good relationship between thrust and time. Further, by making the combustion restricting material soft or having a two-layered or multi-layered structure, the grain 1 will not be destroyed or distorted.
本発明に於いて燃焼制限材がロケツトチヤンバ
ーへ閉じていない場合は、燃焼末期において燃焼
制限材がロケツトチヤンバー４内でグレイン１に
よる支えがなくなり非常に不安定となり、グレイ
ン１の燃焼による高温のため、部分的にちぎれて
ノズル方向へ移動し、ノズルを閉塞する恐れもあ
り、好ましくない。 In the present invention, if the combustion restricting material is not closed to the rocket chamber, the combustion restricting material loses its support from the grains 1 within the rocket chamber 4 at the end of combustion and becomes extremely unstable, resulting in high temperatures caused by the combustion of the grains 1. Therefore, there is a risk that it may partially break off and move toward the nozzle, blocking the nozzle, which is not preferable.
又第１〜２図の如く、端部で閉じた場合は一方
向からの燃焼を設定しているのに対し、閉じない
場合は反対方向からも燃焼が伝わり、燃焼進行距
離は設計した距離より短くなり、所期の性能を確
保し得ない。 Also, as shown in Figures 1 and 2, when the end is closed, combustion is set from one direction, but when it is not closed, combustion is transmitted from the opposite direction as well, and the combustion progress distance is longer than the designed distance. As a result, the desired performance cannot be secured.
本発明は、飛翔体をほぼ垂直に高々度に打ち上
げるのに効果があるのみならず、地面とある角度
をもつた斜方向への飛翔においてもその射程を通
常の内面燃焼形式のロケツトエンジンと比較して
大きく延ばすことを可能とするものである。 The present invention is not only effective in launching a projectile almost vertically to high altitudes, but also in the case of diagonal flight at an angle to the ground. This makes it possible to extend the length significantly.
第１図乃至第３図及び第７図は本発明の一実施
例を示す断面図、第４図は燃焼進行状況を示す要
部拡大図、第５図は第３図の実施例についての推
力Ｆ、時間ｆ曲線のグラフ図、第６図は実際のロ
ケツトに適用した場合のロケツトエンジンの一実
施例の縦断面図である。
Ａ……燃焼進行面、１……グレイン、１ａ……
グレイン燃焼初期面、２……ロケツトエンジン、
３……燃焼制限材、４……ロケツトチヤンバー、
５……ノズル。
Figures 1 to 3 and 7 are cross-sectional views showing one embodiment of the present invention, Figure 4 is an enlarged view of main parts showing the progress of combustion, and Figure 5 is the thrust for the embodiment of Figure 3. F is a graph of the time f curve, and FIG. 6 is a longitudinal sectional view of an embodiment of a rocket engine when applied to an actual rocket. A... Burning progress surface, 1... Grain, 1a...
Grain combustion initial stage, 2...Rocket engine,
3...flammability restriction material, 4...rocket chamber,
5...Nozzle.
Claims (1)
ンが装填され、該グレインはロケツトチヤンバー
の中心軸方向に延びる複数個の板状燃焼制限材で
区画されており、これらの燃焼制限材はロケツト
の中心軸に対し垂直な横断面に於いてロケツトチ
ヤンバーの内周から離隔した円周に沿つて相互に
間隔をおいて配設された円周部分と、これらの円
周部分を夫々ロケツトチヤンバーの内周へ閉じる
連結部分とからなることを特徴とするロケツトエ
ンジン。 ２ 燃焼制限材がロケツトチヤンバーの軸方向に
沿つて部分的に配設されている特許請求の範囲第
１項記載のロケツトエンジン。 ３ 燃焼制限材が軟質材料からなるか或いは二層
又は多層式構造を有する特許請求の範囲第１項又
は第２項記載のロケツトエンジン。[Scope of Claims] 1. A rocket chamber is loaded with internal combustion type grains, and the grains are divided by a plurality of plate-shaped combustion restriction members extending in the direction of the central axis of the rocket chamber, and these combustion restriction members The material has circumferential portions spaced from each other along a circumference spaced apart from the inner circumference of the rocket chamber in a cross section perpendicular to the central axis of the rocket, and these circumferential portions. A rocket engine characterized by comprising a connecting portion that closes to the inner circumference of a rocket chamber. 2. The rocket engine according to claim 1, wherein the combustion restriction material is partially disposed along the axial direction of the rocket chamber. 3. The rocket engine according to claim 1 or 2, wherein the combustion restricting material is made of a soft material or has a two-layer or multi-layer structure.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP188583A JPS59128955A (en) | 1983-01-10 | 1983-01-10 | Rocket engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP188583A JPS59128955A (en) | 1983-01-10 | 1983-01-10 | Rocket engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS59128955A JPS59128955A (en) | 1984-07-25 |
| JPH0319907B2 true JPH0319907B2 (en) | 1991-03-18 |
Family
ID=11514018
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP188583A Granted JPS59128955A (en) | 1983-01-10 | 1983-01-10 | Rocket engine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS59128955A (en) |
Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2742483B1 (en) * | 1995-12-14 | 1998-01-16 | Celerg | SINGLE-COMPOSITION PYROTECHNIC LOADING TO PRODUCE TWO GAS FLOW RATES |
| JP4719182B2 (en) | 2007-05-14 | 2011-07-06 | 三菱重工業株式会社 | 2-pulse rocket motor |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS599744A (en) * | 1982-07-09 | 1984-01-19 | Kokusai Electric Co Ltd | High speed dma (direct memory access) transfer starting circuit |
-
1983
- 1983-01-10 JP JP188583A patent/JPS59128955A/en active Granted
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS599744A (en) * | 1982-07-09 | 1984-01-19 | Kokusai Electric Co Ltd | High speed dma (direct memory access) transfer starting circuit |
Also Published As
| Publication number | Publication date |
|---|---|
| JPS59128955A (en) | 1984-07-25 |
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