JP2749707B2 - Two-stage thrust solid rocket motor - Google Patents
Two-stage thrust solid rocket motorInfo
- Publication number
- JP2749707B2 JP2749707B2 JP14351190A JP14351190A JP2749707B2 JP 2749707 B2 JP2749707 B2 JP 2749707B2 JP 14351190 A JP14351190 A JP 14351190A JP 14351190 A JP14351190 A JP 14351190A JP 2749707 B2 JP2749707 B2 JP 2749707B2
- Authority
- JP
- Japan
- Prior art keywords
- rocket motor
- propellant
- metal
- combustion
- thrust
- 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 - Fee Related
Links
Landscapes
- Shaping Metal By Deep-Drawing, Or The Like (AREA)
- Laminated Bodies (AREA)
Description
【発明の詳細な説明】 〔産業上の利用分野〕 本発明は発進時に大推力を示し、以後低推力を持続す
る一室二層二段推力固体ロケットモータに関する。DETAILED DESCRIPTION OF THE INVENTION [Industrial Application Field] The present invention relates to a single-chamber, two-layer, two-stage thrust solid rocket motor that exhibits a large thrust at the time of starting and maintains a low thrust thereafter.
ロケットモータには発進時の飛翔安定性を高くする目
的で初期（ブースト期）に大推力を発生し、以後低推力
の巡航（サステナ期）を行うために巡航ロケットの外部
にブースタ補助ロケットを付加させる方式のもの、大推
力のブースタロケットモータと小推力のサステナロケッ
トモータを直列に持続する方式のもの等があるが、これ
らは構造が複雑かつ大型となり高価になる。The rocket motor generates a large thrust in the initial stage (boost period) for the purpose of enhancing the flight stability at the start, and then adds a booster auxiliary rocket outside the cruise rocket to perform low-thrust cruise (sustainer period) There are two types: a type in which a booster rocket motor having a large thrust and a type in which a sustainer rocket motor having a small thrust are sustained in series. However, these are complicated, large-sized, and expensive.
一方、ロケットモータの燃焼室を一室とし、内部の内
面燃焼推進薬を二層とし、内層に高燃速推進薬、外層に
低燃速推進薬を配した二段推力固体ロケットモータは単
純、小型で安価であるが、高燃速推進薬は固体酸化剤の
粒径が細かく、かつその含有量が多く、更に高燃速触媒
を多量に含むために機械的強度が低下し、低温では内面
に大きな引張応力を発生するので、推進薬に致命的な亀
裂欠陥を生じる。また、製造工程も初めに内層になる高
燃速推進薬の円筒を予め硬化成形しておき、これをロケ
ットモータ燃焼室の中央部に配置して、その外側と燃焼
室内壁との間に外層となる低燃速推進薬を注入した後、
硬化させ成形するというように製造工程が２倍となる。
更に内層と外層間の推進薬が剥離を起こす危険があり、
両層間の接着力を維持させるには高度の技術を必要とす
る。On the other hand, a two-stage thrust solid rocket motor with a single rocket motor combustion chamber, two layers of internal combustion propellant inside, a high fuel propellant in the inner layer, and a low fuel propellant in the outer layer is simple, Although small in size and inexpensive, high-fuel propellants have a small solid oxidizer particle size, a high content, and a large amount of high-fuel catalyst, resulting in low mechanical strength. It generates a large tensile stress in the propellant, causing fatal crack defects in the propellant. In the manufacturing process, a cylinder of the high-fuel-speed propellant, which is to be the inner layer, is first hardened and formed in advance, and this is disposed in the center of the rocket motor combustion chamber, and the outer layer is interposed between the outside and the combustion chamber wall. After injecting a low-fuel propellant,
The manufacturing process is doubled, such as curing and molding.
Furthermore, there is a risk that the propellant between the inner layer and the outer layer may peel off,
A high level of technology is required to maintain the adhesion between the two layers.
本発明の目的は一段推力ロケットモータと殆ど同一の
製造工程で、信頼性の高い二段推力ロケットモータを得
ることにある。An object of the present invention is to obtain a highly reliable two-stage thrust rocket motor in almost the same manufacturing process as a one-stage thrust rocket motor.
本発明者らは、上記の目的を達成すべく鋭意検討した
結果、本発明を完成するに到った。The present inventors have conducted intensive studies to achieve the above object, and as a result, completed the present invention.
即ち、本発明は一室二層二段推力固体ロケットモータ
において、内面燃焼方式推進薬の内層として円筒状で金
属製の通気性発泡体を配置したことを特徴とする二段推
力固体ロケットモータを提供するものである。That is, the present invention provides a single-chamber, two-layer, two-stage thrust solid rocket motor, wherein a cylindrical, metal breathable foam is disposed as an inner layer of an inner combustion type propellant, To provide.
本発明において内面燃焼方式推進薬として用いる推進
薬は一種類でも良く、高燃速推進薬をあえて用いる必要
はなく燃焼室の中央に推進薬の内層として円筒状で金属
製の通気性発泡体を配置するのみで二段推力を得ること
ができ、かつ低温に曝された場合も推進薬内面に亀裂を
生じることがなく、耐環境性も向上する。勿論、ブース
タ期とサステナ期の推力比をより大きくする必要がある
場合は内面燃焼方式推進薬の内層推進薬として高燃速推
進薬を使用することもでき、この場合は内層に配置され
た前記発泡体が低温において推進薬内面に低温で亀裂を
生じることを防ぐ効果もある。In the present invention, the propellant used as the internal combustion type propellant may be one kind, and it is not necessary to use a high fuel speed propellant, and a cylindrical metal air-permeable foam is used as the inner layer of the propellant in the center of the combustion chamber. A two-stage thrust can be obtained only by arranging, and even when exposed to a low temperature, cracks do not occur on the inner surface of the propellant and environmental resistance is improved. Of course, if it is necessary to increase the thrust ratio between the booster period and the sustainer period, a high fuel speed propellant can be used as the inner layer propellant of the inner combustion type propellant, and in this case, the propellant disposed in the inner layer It also has the effect of preventing the foam from cracking at low temperatures at the propellant inner surface at low temperatures.
本発明の詳細を図面により説明すると、第１図は本発
明になる一室二層二段推力固体ロケットモータの断面略
示図であり、第２図は同じく胴部直角方向の断面略示図
である。第１図において１は推進薬を、２は金属通気性
発泡体を、３は点火器、４はロケットモータ燃焼室、５
はレストリクタを示す。FIG. 1 is a schematic cross-sectional view of a single-chamber, two-layer, two-stage thrust solid rocket motor according to the present invention, and FIG. It is. In FIG. 1, 1 is a propellant, 2 is a metal permeable foam, 3 is an igniter, 4 is a rocket motor combustion chamber, 5
Indicates a restrictor.
本発明になるロケットモータの製造方法を説明する
と、予め第３図に示すような横断面の中央部に光芒形の
空間部を有する金属製の通気性発泡体６を制作し、その
光芒形空間部７に第４図に示すアルミニウム等の金属か
らなる同じ断面形をもつ金属型８をはめ込み、推進薬注
型芯とする。使用する金属発泡体の通気性の一例を示す
と、厚さ1cm、平均孔径3.2mm、空気流速2m/secで圧力損
失は約7mmHgのものである。孔径が大きくなるに従い、
圧力損失は低下し通気性が向上し、推進薬の注型性が向
上する。発泡体を構成する金属としてはアルミニウム、
銀が好ましく、アルミニウムの材質は純アルミニウムの
他に、A6101、A1070、AC2A、AC4CH等のアルミニウム合
金も使用できる。金属発泡体の見掛密度は0.05g/cm3程
度であり、熱伝導率は重量比で銅の約２倍である。ブー
スタ期とサステナ期の高い推力比は金属気泡壁の厚さを
薄くするか、より熱伝導率の高い銀を使用することで得
られる。The method of manufacturing the rocket motor according to the present invention will be described. A metal breathable foam 6 having a light beam-shaped space at the center of the cross section as shown in FIG. A metal mold 8 having the same cross-sectional shape made of a metal such as aluminum shown in FIG. 4 is fitted into the part 7 to obtain a propellant injection core. As an example of the air permeability of the metal foam used, the thickness is 1 cm, the average pore diameter is 3.2 mm, the air flow rate is 2 m / sec, and the pressure loss is about 7 mmHg. As the pore size increases,
Pressure loss is reduced, air permeability is improved, and propellant casting is improved. Aluminum as the metal that constitutes the foam,
Silver is preferred, and aluminum materials such as aluminum alloys such as A6101, A1070, AC2A, AC4CH can be used in addition to pure aluminum. The apparent density of the metal foam is about 0.05 g / cm 3 , and the thermal conductivity is about twice that of copper by weight. High thrust ratios in the booster and sustainer phases can be obtained by reducing the thickness of the metal bubble walls or by using silver with higher thermal conductivity.
上記の注型芯を第５図中のロケットモータ燃焼室９の
中心に設置し、通常の手段で推進薬１のスラリーを注型
し、加熱硬化した後、アルミニウム金属型８を抜芯し、
端部をレストリクタ５仕上げし、製造を完了する。The above casting core was placed at the center of the rocket motor combustion chamber 9 in FIG. 5, the slurry of the propellant 1 was cast by ordinary means, and after being cured by heating, the aluminum metal mold 8 was cored.
The end is finished with a restrictor 5 to complete the manufacture.
この様にして製造したロケットモータを通常の手法で
点火信号によって燃焼させると、点火器の火炎が推進薬
内面を着火させるが、推進薬内面から内層部には熱伝導
率の高いアルミニウム、銀等の金属通気性発泡体が存在
しており、この発泡体壁の金属薄箔は燃焼部からの熱を
推進薬の燃焼進行方向に急速に伝えるために金属発泡体
壁に接している推進薬は温度が上昇し、燃焼速度が早く
なり、結果的には金属発泡体を含む推進薬全体の燃焼速
度が金属発泡体を含まぬものより数倍早くなる。When the rocket motor manufactured in this way is burned by an ignition signal in a usual manner, the flame of the igniter ignites the inner surface of the propellant, but aluminum, silver, etc. having high thermal conductivity are formed from the inner surface of the propellant to the inner layer. There is a metal-permeable foam of the following, and the thin metal foil of the foam wall is in contact with the metal foam wall in order to rapidly transfer the heat from the combustion part in the direction of combustion of the propellant. The temperature rises and the burning rate is faster, resulting in a burning rate of the entire propellant containing the metal foam that is several times faster than that without the metal foam.
このように金属の熱伝導を利用して推進薬の燃焼速度
を上げる方式は従来から銀線やハネカムを用いることが
提案されているが、端面燃焼のみにしか適用できなかっ
たり、又は材料が軟弱なために一定の形状を保つことが
困難であった。また金属箔片を推進薬中に予め練込む方
式は注形に問題が多く、再現性に乏しかった。As described above, the method of increasing the burning speed of the propellant using the heat conduction of the metal has been proposed to use a silver wire or a honeycomb, but it can be applied only to the end face combustion, or the material is soft. Therefore, it was difficult to maintain a certain shape. In addition, the method of kneading the metal foil pieces into the propellant beforehand has many problems in casting and has poor reproducibility.
一方、本発明で用いる金属発泡体は各気孔が多角形で
あるために、３次元的に等方向性であり、かつ剛性が高
いので任意の形状に加工することができ、内面燃焼方式
の他に端面燃焼方式にも適用できる。On the other hand, the metal foam used in the present invention is three-dimensionally isotropic since the pores are polygonal, and can be processed into an arbitrary shape because of high rigidity. It can also be applied to the end face combustion method.
以下実施例にて本発明を説明するが、本発明はこれら
の実施例に限定されるものではない。Hereinafter, the present invention will be described with reference to examples, but the present invention is not limited to these examples.
第３図に示す如き中央部に光芒形の空間部を有する外
径が6cm、長さ20cmの円筒形のアルミニウム金属通気性
発泡体（平均孔径6.4mm）の光芒形空間部７に、第４図
に示すアルミニウム金属型８をはめ込み、注型芯とし、
内径8cm、平行部長さが20cmの円筒形のロケットモータ
燃焼室の中心部に設置し、燃焼速度7mm/secを示す粘度
２キロポイズの過塩素酸アンモニウム−ポリブタジエン
系推進薬スラリーを注型し65℃で３日間加熱硬化した
後、中心部のアルミニウム金属型を抜芯した。両端部を
耐熱材でレストリクタ加工し、本発明のロケットモータ
の製造を完了した。As shown in FIG. 3, a light beam type space portion 7 of a cylindrical aluminum metal breathable foam (average pore diameter 6.4 mm) having an outer diameter of 6 cm and a length of 20 cm having a light beam type space portion at the center portion, Insert the aluminum metal mold 8 shown in the figure to make a casting core,
Installed in the center of a cylindrical rocket motor combustion chamber with an inner diameter of 8 cm and a parallel part length of 20 cm, cast an ammonium perchlorate-polybutadiene-based propellant slurry having a viscosity of 2 kpoise and a combustion speed of 7 mm / sec. After heating and curing for 3 days, the aluminum metal mold at the center was cored. Both ends were subjected to restrictor processing with a heat-resistant material to complete the manufacture of the rocket motor of the present invention.
比較例として燃焼速度25mm/secを示す粘度12キロポイ
ズの過塩素酸アンモニウム−ポリブタジエン系推進薬
で、中央部に光芒形の空間部を有する外径が6cm、長さ2
0cmの円筒形のブースタ推進薬を予め製造しておき、ア
ルミニウム金属型が付いたままこれを注型芯として、内
径8cm、平行部長さが20cmの円筒形のロケットモータの
中心部に設置し、注型芯の外側と燃焼室内壁との間にサ
ステナ推進薬として燃焼速度7mm/secを示す粘度２キロ
ポイズの過塩素酸アンモニウム−ポリブタジエン系推進
薬スラリーを注型し、65℃で３日間加熱硬化した後、抜
芯し、両端部を耐熱材でレストリクタ加工して比較用ロ
ケットモータを作製した。As a comparative example, an ammonium perchlorate having a viscosity of 12 kpoise and a polybutadiene-based propellant exhibiting a combustion speed of 25 mm / sec, an outer diameter having a light-shaped space in the center is 6 cm, and the length is 2
A cylindrical booster propellant of 0 cm is manufactured in advance, and this is used as a casting core with the aluminum metal mold attached, installed at the center of a cylindrical rocket motor having an inner diameter of 8 cm and a parallel part length of 20 cm, A 2 kilopoise ammonium perchlorate-polybutadiene-based propellant slurry exhibiting a burning speed of 7 mm / sec is cast as a sustaining propellant between the outside of the casting core and the combustion chamber wall, and is heat-cured at 65 ° C for 3 days. Then, the core was removed, and both ends were subjected to restrictor processing with a heat-resistant material to produce a rocket motor for comparison.
本発明のロケットモータと比較用ロケットモータを−
60℃に調温した後、−60℃において振動試験を行うと、
比較ロケットモータは内面光芒形の空間部の底部にモー
タ軸に添って亀裂が発生しており爆発の危険があり、燃
焼試験に供し得なかった。The rocket motor of the present invention and the rocket motor for comparison
After adjusting the temperature to 60 ° C, perform a vibration test at -60 ° C.
The comparative rocket motor had a crack along the motor shaft at the bottom of the inner beam-shaped space, and there was a danger of explosion, so it could not be subjected to a combustion test.
本発明のロケットモータは内面目視及びＸ線による非
破壊検査に於いても全く異常がないので、点火器を取り
つけ−60℃に再び調温した後、燃焼スタンドに取りつけ
燃焼試験を行うと、ブースタ期に292kgの推力を0.61秒
間発生し、サステナ期に97.4kgの推力を1.83秒間持続
し、正常な二段推力の燃焼を示した。Since the rocket motor of the present invention has no abnormality in the internal visual inspection and non-destructive inspection by X-ray, the igniter is mounted, the temperature is adjusted again to -60 ° C, and the rocket motor is mounted on the combustion stand and the combustion test is performed. A thrust of 292 kg was generated for 0.61 seconds during the period, and a thrust of 97.4 kg was sustained for 1.83 seconds during the sustainer period, showing normal two-stage thrust combustion.
実施例で示すように従来の二段推力固体ロケットモー
タは製造工程が複数回であり、かつ低温環境では強度不
足から推進薬内面に亀裂を発生し使用に耐えない。As shown in the embodiment, the conventional two-stage thrust solid rocket motor has a plurality of manufacturing steps, and in a low-temperature environment, cracks are generated on the inner surface of the propellant due to insufficient strength and cannot be used.
これに対し、本発明のロケットモータは推進薬の注型
工程が一回ですむため製造が簡単になり、かつ−60℃の
低温環境にも耐え、正常な燃焼を示し、優れた性能を発
揮する。On the other hand, the rocket motor of the present invention requires only one propellant injection step, which simplifies manufacturing, withstands low-temperature environments of -60 ° C, shows normal combustion, and exhibits excellent performance. I do.
第１図は本発明のロケットモータの断面略示図、第２図
は本発明のロケットモータの胴部直角方向の断面略示
図、第３図は本発明に用いられる金属発泡体の斜視図、
第４図は金属型の斜視図、第５図は製造工程中のロケッ
トモータを示す断面略示図である。 １……推進薬 2,6……金属発泡体 ３……点火器 4,9……燃焼室 ５……レストリクタ ７……光芒形空間部 ８……金属型1 is a schematic cross-sectional view of a rocket motor of the present invention, FIG. 2 is a schematic cross-sectional view of a rocket motor of the present invention in a direction perpendicular to the trunk, and FIG. 3 is a perspective view of a metal foam used in the present invention. ,
FIG. 4 is a perspective view of a metal mold, and FIG. 5 is a schematic sectional view showing a rocket motor during a manufacturing process. 1 ... Propellant 2,6 ... Metal foam 3 ... Ignition device 4,9 ... Combustion chamber 5 ... Restrictor 7 ... Light beam type space 8 ... Metal type
Claims (2)
いて、内面燃焼方式推進薬の内層として円筒状で金属製
の通気性発泡体を配置したことを特徴とする二段推力固
体ロケットモータ。1. A single-stage, two-layer, two-stage thrust solid rocket motor, wherein a cylindrical, metallic, breathable foam is disposed as an inner layer of an internal combustion type propellant.
銀製である請求項１記載の二段推力固体ロケットモー
タ。2. The two-stage thrust solid rocket motor according to claim 1, wherein the metal breathable foam is made of aluminum or silver.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP14351190A JP2749707B2 (en) | 1990-06-01 | 1990-06-01 | Two-stage thrust solid rocket motor |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP14351190A JP2749707B2 (en) | 1990-06-01 | 1990-06-01 | Two-stage thrust solid rocket motor |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH0441964A JPH0441964A (en) | 1992-02-12 |
| JP2749707B2 true JP2749707B2 (en) | 1998-05-13 |
Family
ID=15340437
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP14351190A Expired - Fee Related JP2749707B2 (en) | 1990-06-01 | 1990-06-01 | Two-stage thrust solid rocket motor |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP2749707B2 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110566366A (en) * | 2019-09-02 | 2019-12-13 | 湖北三江航天江河化工科技有限公司 | Combined core mold structure for rocket engine grain molding and use method thereof |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP4719182B2 (en) | 2007-05-14 | 2011-07-06 | 三菱重工業株式会社 | 2-pulse rocket motor |
| JP4919980B2 (en) * | 2008-01-28 | 2012-04-18 | 株式会社Ｉｈｉエアロスペース | End face combustion type gas generator |
| JP5303369B2 (en) * | 2009-06-09 | 2013-10-02 | 株式会社Ｉｈｉエアロスペース | Solid rocket motor propellant molding method |
-
1990
- 1990-06-01 JP JP14351190A patent/JP2749707B2/en not_active Expired - Fee Related
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110566366A (en) * | 2019-09-02 | 2019-12-13 | 湖北三江航天江河化工科技有限公司 | Combined core mold structure for rocket engine grain molding and use method thereof |
| CN110566366B (en) * | 2019-09-02 | 2021-07-13 | 湖北三江航天江河化工科技有限公司 | Combined core mold structure for rocket engine grain molding and use method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| JPH0441964A (en) | 1992-02-12 |
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