JPH06330816A - High temperature combustion engine - Google Patents
High temperature combustion engineInfo
- Publication number
- JPH06330816A JPH06330816A JP11696193A JP11696193A JPH06330816A JP H06330816 A JPH06330816 A JP H06330816A JP 11696193 A JP11696193 A JP 11696193A JP 11696193 A JP11696193 A JP 11696193A JP H06330816 A JPH06330816 A JP H06330816A
- Authority
- JP
- Japan
- Prior art keywords
- combustion chamber
- cooling
- cooling passage
- rocket engine
- engine
- 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.)
- Granted
Links
Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は、燃焼室を形成する壁面
内に冷却媒体が流される冷却通路をもつロケットエンジ
ン、スクラムジェットエンジンなどの高温燃焼エンジン
に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a high temperature combustion engine such as a rocket engine or a scramjet engine having a cooling passage through which a cooling medium flows in a wall forming a combustion chamber.
【０００２】[0002]
【従来の技術】燃焼室に冷却通路をもつロケットエンジ
ンでは、冷却通路内を流れる冷却媒体が伝熱面上から気
泡の発生、成長、離脱を繰り返えす、いわゆる核沸騰の
段階で伝熱面から気泡をすみやかに離脱させることによ
って、伝熱面が蒸気膜で覆われる、いわゆる膜沸騰を起
こし難くするよう冷却媒体の流量を増やし流速を上げて
いた。2. Description of the Related Art In a rocket engine having a cooling passage in a combustion chamber, a cooling medium flowing in the cooling passage repeats the generation, growth and separation of bubbles from the heat transfer surface, that is, at the so-called nucleate boiling stage. By rapidly removing the bubbles from the heat transfer surface, the flow rate of the cooling medium was increased so that the heat transfer surface was covered with the vapor film, so-called film boiling was less likely to occur.
【０００３】或いは、冷却通路内で核沸騰そのものを起
こし難くするため冷却通路の高温側の壁に耐熱コーテン
グを施したりしていた。Alternatively, in order to make nucleate boiling itself less likely to occur in the cooling passage, heat-resistant coating has been applied to the high temperature side wall of the cooling passage.
【０００４】[0004]
【発明が解決しようとする課題】ロケットエンジン燃焼
室の冷却通路内で、冷却媒体に気液２相が混在する臨界
圧以下で冷却させるには、冷却媒体の吸収する熱流束が
冷却媒体の沸騰を起こす熱流束（以下バーンアウト熱流
束という）を越えないようにする必要がある。In order to cool the cooling medium in the cooling passage of the rocket engine combustion chamber at a critical pressure below which the gas-liquid two phases coexist in the cooling medium, the heat flux absorbed by the cooling medium is boiled. It is necessary not to exceed the heat flux (hereinafter referred to as burnout heat flux) that causes
【０００５】その為には、熱伝達率等とは無関係に、バ
ーンアウト熱流束を大きくし、冷却媒体の吸収熱流束を
それ以下にすることが重要である。一般にバーンアウト
熱流束は冷却媒体の流速のべき乗に比例するので、バー
ンアウト熱流束を大きくして冷却媒体が沸騰しにくくす
るためには、冷却媒体の流速を増さなければならない。For that purpose, it is important to increase the burnout heat flux and make the absorption heat flux of the cooling medium lower than that regardless of the heat transfer coefficient and the like. Generally, the burnout heat flux is proportional to the power of the flow velocity of the cooling medium. Therefore, in order to increase the burnout heat flux to prevent the cooling medium from boiling, the flow velocity of the cooling medium must be increased.
【０００６】しかしながら、圧力損失は、流速の２乗に
比例するので流速の増加は圧力損失につながり、最終的
には、タンク圧やポンプ圧の増加につながり、それだけ
重量が増すのでエンジンの正味の性能が低下することに
なる。However, since the pressure loss is proportional to the square of the flow velocity, an increase in the flow velocity leads to a pressure loss, and eventually to an increase in tank pressure or pump pressure, which increases the weight and thus the net weight of the engine. Performance will be reduced.
【０００７】従って、エンジンで燃焼させる燃料を冷却
媒体として燃焼室の冷却通路に流すようにした、いわゆ
る再生冷却燃焼室をもつロケットエンジンでは大きな問
題を生ずる。また、冷却通路に核沸騰防止のための断熱
コーティングを施しているものは、コーティングがはが
れやすいため、再使用型のロケットエンジンには向かな
い。Therefore, a big problem occurs in a rocket engine having a so-called regenerative cooling combustion chamber in which fuel to be burned in the engine is made to flow as a cooling medium into the cooling passage of the combustion chamber. In addition, a cooling passage having a heat insulating coating for preventing nucleate boiling is not suitable for a reusable rocket engine because the coating is easily peeled off.
【０００８】本発明は、冷却媒体の流れに対する圧損が
少く、耐熱性能を高めた、燃焼室を形成する壁面に冷却
通路をもつ燃焼室を具えた高温燃焼エンジンを提供する
ことを課題としている。It is an object of the present invention to provide a high temperature combustion engine having a combustion chamber having a cooling passage on the wall surface forming the combustion chamber, which has a small pressure loss with respect to the flow of the cooling medium and has improved heat resistance.
【０００９】[0009]
【課題を解決するための手段】本発明では、燃焼室を形
成する壁面内に冷却媒体が流される冷却通路をもつ高温
燃焼エンジンにおける前記した課題を解決するため前記
した冷却通路内に２つのテーパ面をもち、冷却媒体の流
れ方向に伸びる少くとも１つのフィンを配設する。According to the present invention, in order to solve the above-mentioned problems in a high temperature combustion engine having a cooling passage through which a cooling medium flows in a wall surface forming a combustion chamber, two tapers are provided in the cooling passage. Arranged is at least one fin having a face and extending in the direction of flow of the cooling medium.
【００１０】また、本発明では、燃焼室を形成する壁面
内に冷却媒体として燃焼前の燃料が流される冷却通路を
もつロケットエンジンにおいて、前記した課題を解決す
るために、前記冷却通路を燃焼室後部の入口マニホール
ドから燃焼室前部の出口マニホールドへ燃焼室の軸方向
に伸びた複数本の冷却通路によって構成し、同冷却通路
内の各々に２つのテーパ面をもつ少くとも１つのフィン
を燃焼室軸方向に沿って配設した構成を採用する。Further, according to the present invention, in a rocket engine having a cooling passage through which fuel before combustion flows as a cooling medium in a wall surface forming a combustion chamber, in order to solve the above-mentioned problems, the cooling passage is provided in the combustion chamber. Composed of multiple cooling passages extending in the axial direction of the combustion chamber from the rear inlet manifold to the front outlet manifold, and burns at least one fin with two tapered surfaces in each cooling passage. The structure is arranged along the chamber axis.
【００１１】[0011]
【作用】一般にロケットエンジンなどでは、冷却通路内
で冷却媒体が核沸騰から膜沸騰へと沸騰が進行し、つい
に燃焼室が溶融するに至る。そこで、冷却通路内にフィ
ンを配設し、伝熱面に平坦部を無くすることで前述のバ
ーンアウト熱流束を飛躍的に高めることができる。In general, in a rocket engine or the like, the cooling medium boils in the cooling passage from nucleate boiling to film boiling, and finally the combustion chamber melts. Therefore, by disposing fins in the cooling passage and eliminating the flat portion on the heat transfer surface, the above-mentioned burnout heat flux can be dramatically increased.
【００１２】本発明はこの考えに基づき、燃焼室を形成
する壁面の冷却通路に２つのテーパ面をもち冷却媒体の
流れ方向に伸びる１個又は２個以上のフィンを設けるも
ので、これによると、冷却媒体の沸騰により伝熱面に発
生する気泡がフィンを形成する２つのテーパ面に沿って
速やかに離脱され、伝熱面での膜沸騰が起こり難くな
る。Based on this idea, the present invention provides one or more fins having two tapered surfaces and extending in the flow direction of the cooling medium in the cooling passage of the wall surface forming the combustion chamber. The bubbles generated on the heat transfer surface due to the boiling of the cooling medium are quickly separated along the two tapered surfaces forming the fins, and the film boiling on the heat transfer surface is less likely to occur.
【００１３】また、本発明によってロケットエンジンに
前記した構成を採用することによって燃焼室の耐熱特性
が大きく向上し再使用を可能とする。Further, by adopting the above-mentioned structure to the rocket engine according to the present invention, the heat resistance of the combustion chamber is greatly improved and the combustion chamber can be reused.
【００１４】[0014]
【実施例】以下、本発明を図示した実施例に基いて具体
的に説明する。なお、以下の実施例は、エンジンで燃焼
する燃料で燃焼室を冷却するようにした再生冷却式ロケ
ットエンジンに対し本発明を適用した場合である。図１
において、Ｃは燃焼室であり内筒１と外筒２によって構
成されている。EXAMPLES The present invention will be specifically described below based on illustrated examples. In addition, the following embodiment is a case where the present invention is applied to a regenerative cooling rocket engine in which a combustion chamber is cooled by fuel burned in the engine. Figure 1
In the figure, C is a combustion chamber, which is composed of an inner cylinder 1 and an outer cylinder 2.
【００１５】３は、燃焼室Ｃの後部の周囲に設けられた
冷却媒体のための入口マニホールドであり、４は燃焼室
Ｃの前部の周囲に設けられた冷却媒体のための出口マニ
ホールドである。内筒１には、図２の断面図に拡大して
示されているように燃焼室Ｃの軸方向に入口マニホール
ド３から出口マニホールド４に互いに平行に伸びている
複数個の冷却通路５が形成されている。Reference numeral 3 is an inlet manifold for the cooling medium provided around the rear portion of the combustion chamber C, and 4 is an outlet manifold for the cooling medium provided around the front portion of the combustion chamber C. . A plurality of cooling passages 5 extending in the axial direction of the combustion chamber C from the inlet manifold 3 to the outlet manifold 4 in parallel to each other are formed in the inner cylinder 1 as shown in an enlarged view in the cross-sectional view of FIG. Has been done.
【００１６】各冷却通路５内には、頂部が６０°の角度
をなす２つのテーパ面６をもつフィン７が冷却通路５に
沿って伸びて配設されている。燃焼室Ｃで燃焼されるモ
ノメチルヒドラジンなどの燃料が入口配管８から入口マ
ニホールド３を経て内筒１内の複数個の冷却通路５へ分
配して流され、燃焼室Ｃを冷却して温度上昇し高エネル
ギー状態となって出口マニホールド４で集められて出口
配管９から流出する。Within each cooling passage 5, fins 7 having two tapered surfaces 6 whose tops form an angle of 60 ° are arranged so as to extend along the cooling passage 5. Fuel such as monomethylhydrazine that is burned in the combustion chamber C is distributed from the inlet pipe 8 through the inlet manifold 3 to the plurality of cooling passages 5 in the inner cylinder 1 to cool the combustion chamber C and raise its temperature. The high energy state is collected in the outlet manifold 4 and flows out from the outlet pipe 9.
【００１７】燃料は、出口配管９から噴射器入口１０を
経て燃焼室Ｃの燃料噴射器へ導かれ、酸化剤入口１１か
ら供給される酸化剤によって燃焼室Ｃ内で燃焼される。
なお、図においてＰ1 ，Ｐ2 ，Ｐ3 ，Ｐ4 は、それぞ
れ、圧力計測ポートを示している。The fuel is guided from the outlet pipe 9 to the fuel injector of the combustion chamber C via the injector inlet 10, and is burned in the combustion chamber C by the oxidant supplied from the oxidant inlet 11.
In the figure, P 1 , P 2 , P 3 , and P 4 respectively indicate pressure measurement ports.
【００１８】次に図３は、図１に示した燃焼室Ｃの製造
課程を示している。図３の工程（１）はアルミマンドレ
ルで燃焼室の内面形状を形成した状態を示している。工
程（２）は、工程（１）でつくられたアルミマンドレル
の上に電鋳によって内筒１を形成した状態を示してい
る。工程（３）は、工程（２）でつくられた電鋳層の外
面上にその長さ方向に伸びる複数本の冷却通路５とフィ
ン７を機械加工によって形成した状態を示している。Next, FIG. 3 shows a manufacturing process of the combustion chamber C shown in FIG. Step (1) in FIG. 3 shows a state in which the inner surface shape of the combustion chamber is formed by an aluminum mandrel. Step (2) shows a state in which the inner cylinder 1 is formed by electroforming on the aluminum mandrel made in step (1). Step (3) shows a state in which a plurality of cooling passages 5 and fins 7 extending in the lengthwise direction are formed by machining on the outer surface of the electroformed layer produced in step (2).
【００１９】電鋳層の外面に冷却通路５とフィン７を機
械加工するには図４〜図６に示すようなカッターを使
う。図４と図６に示すカッターは、それぞれ（ｃ）の図
に加工溝の断面が示されているように１つの山形のフィ
ンをもつ冷却通路を加工する場合に使うもの、図５に示
すカッターは（ｃ）の図に示す加工溝のように４本のフ
ィンをもつ冷却通路を加工する場合のカッターの例であ
る。To machine the cooling passages 5 and the fins 7 on the outer surface of the electroformed layer, a cutter as shown in FIGS. 4 to 6 is used. The cutters shown in FIGS. 4 and 6 are used for machining a cooling passage having one chevron fin as shown in the sectional view of the machining groove in (c), respectively, and the cutter shown in FIG. Is an example of a cutter for machining a cooling passage having four fins like the machining groove shown in FIG.
【００２０】形成するフィンの先端は、冷却通路内で沸
騰が起きないようにｃ０〜ｃ０．１（chamfer 数）が必
要であり、平坦部を極力つくらないようにする。放電加
工では角が落ちて了うので前記したように機械加工で形
成する。The tip of the fin to be formed needs to have c0 to c0.1 (chamfer number) so that boiling does not occur in the cooling passage, and the flat portion is not formed as much as possible. In the electric discharge machining, since the corners are finished, it is formed by machining as described above.
【００２１】また、膜沸騰を抑えるためには図５に示し
たように冷却通路の伝熱面全体にフィンを作り平行部を
設けないのが最も望ましいが、カッターによる加工が困
難となり費用もかかるのでフィンの数を何個にするかは
対象に応じて決定する。Further, in order to suppress film boiling, it is most desirable to form fins on the entire heat transfer surface of the cooling passage and not to provide a parallel portion as shown in FIG. 5, but it is difficult and expensive to process with a cutter. Therefore, the number of fins is determined according to the object.
【００２２】なお、本発明を再成冷却燃焼室に適用する
場合は、推薬を臨界圧以下で扱うエンジンに対してであ
り、一般にその場合にはノズルスカートは輻射冷却とな
るため従来技術でつくられる。また、前述の実施例で
は、２つのテーパ面の交わる角度が６０°の場合を説明
したが、３０°，４５°，９０°など適宜選択してよ
い。When the present invention is applied to a regenerative cooling combustion chamber, it is applied to an engine in which propellant is handled at a critical pressure or less. Generally, in that case, since the nozzle skirt is radiatively cooled, the conventional technique is used. able to make. Further, in the above-described embodiment, the case where the angle at which the two tapered surfaces intersect is 60 ° has been described, but it may be appropriately selected such as 30 °, 45 °, 90 °.
【００２３】ここで再び図３に戻って、工程（３）にお
いて冷却通路５とフィン７を形成したあと、溝にはワッ
クスが充填され工程（２）と同様の外形に戻される。工
程（４）は、工程（３）の後、その表面上に重ねて電鋳
を行ない外側の内筒１を形成した状態を示している。工
程（５）は、前記したように冷却通路５とフィン７を内
部にもつ内筒１をつくったのちアルミマンドレルを溶解
除去して内筒を得た状態を示している。こうして得た内
筒１の外面上に所要の外筒２を形成して燃焼室Ｃがつく
られる。Now, returning to FIG. 3 again, after the cooling passages 5 and the fins 7 are formed in the step (3), the groove is filled with wax and the outer shape similar to that in the step (2) is restored. Step (4) shows a state in which after the step (3), the outer inner cylinder 1 is formed on the surface by electroforming. Step (5) shows a state in which the inner cylinder 1 having the cooling passages 5 and the fins 7 therein is formed as described above, and then the aluminum mandrel is dissolved and removed to obtain the inner cylinder. A combustion chamber C is created by forming the required outer cylinder 2 on the outer surface of the inner cylinder 1 thus obtained.
【００２４】[0024]
【発明の効果】以上、具体的に説明したように、本発明
により燃焼室に２つのテーパ面からなるフィンをもつ冷
却通路を設けることにより、冷却通路において冷却媒体
が発する気泡の離脱をスムーズにし耐熱特性を高め圧損
を大きくすることなく燃焼室の冷却を容易にすることを
可能にした。As described above in detail, according to the present invention, by providing the cooling passage having the fins having the two tapered surfaces in the combustion chamber, it is possible to smoothly remove the bubbles generated by the cooling medium in the cooling passage. It has improved heat resistance and facilitated cooling of the combustion chamber without increasing pressure loss.
【図１】本発明をロケットエンジンに適用した実施例を
示す断面図、FIG. 1 is a sectional view showing an embodiment in which the present invention is applied to a rocket engine,
【図２】図１のＡ−Ａ線に沿う拡大断面図、FIG. 2 is an enlarged cross-sectional view taken along the line AA of FIG.
【図３】図１に示す燃焼室を製造する工程を示した図面
で、（１）〜（４）において、（ａ）は斜視図、（ｂ）
は断面図、FIG. 3 is a drawing showing a process of manufacturing the combustion chamber shown in FIG. 1, wherein (a) is a perspective view and (b) is (1) to (4).
Is a sectional view,
【図４】本発明による冷却通路を加工するためのカッタ
ーを示す図面で（ａ）は断面図、（ｂ）は正面図、
（ｃ）は加工溝の断面を示す。4A and 4B are views showing a cutter for processing a cooling passage according to the present invention, in which FIG. 4A is a sectional view and FIG.
(C) shows the cross section of a processed groove.
【図５】他のカッターを示す図４と同様の図面、5 is a drawing similar to FIG. 4 showing another cutter,
【図６】更に他のカッターを示す図４と同様の図面。FIG. 6 is a view similar to FIG. 4 showing still another cutter.
１ 内筒 ２ 外筒 ３ 入口マニホールド ４ 出口マニホールド ５ 冷却通路 ６ テーパ面 ７ フィン 1 Inner Cylinder 2 Outer Cylinder 3 Inlet Manifold 4 Outlet Manifold 5 Cooling Passage 6 Tapered Surface 7 Fins
───────────────────────────────────────────────────── フロントページの続き (72)発明者 外川 守人 愛知県小牧市大字東田中1200番地 三菱重 工業株式会社名古屋誘導推進システム製作 所内 (72)発明者 三木 陽一郎 愛知県小牧市大字東田中1200番地 三菱重 工業株式会社名古屋誘導推進システム製作 所内 ─────────────────────────────────────────────────── ─── Continuation of front page (72) Inventor Morito Sotokawa 1200, Higashi-Tanaka, Komaki-shi, Aichi Mitsubishi Heavy Industries, Ltd. Nagoya guidance propulsion system manufacturing plant (72) Inventor Yoichiro Miki 1200, Higashi-Tanaka, Komaki, Aichi Mitsubishi Heavy Industry Co., Ltd. Nagoya guidance propulsion system manufacturing plant
Claims (2)
される冷却通路をもつ高温燃焼エンジンにおいて、前記
冷却通路内に、２つのテーパ面をもち前記冷却媒体の流
れ方向に伸びる少くとも１つのフィンを配設したことを
特徴とする高温燃焼エンジン。1. A high temperature combustion engine having a cooling passage through which a cooling medium flows in a wall forming a combustion chamber, wherein the cooling passage has at least one tapered surface extending in the flow direction of the cooling medium. A high-temperature combustion engine that features two fins.
て燃焼前の燃料が流される冷却通路をもつロケットエン
ジンにおいて、前記冷却通路が前記燃焼室後部の入口マ
ニホールドから燃焼室前部の出口マニホールドへ前記燃
焼室の軸方向に伸びる複数本の冷却通路によって構成さ
れ、同冷却通路内の各々に２つのテーパ面をもつ少くと
も１つのフィンが前記軸方向に沿って配設されているこ
とを特徴とするロケットエンジン。2. A rocket engine having a cooling passage through which fuel before combustion flows as a cooling medium in a wall surface forming a combustion chamber, wherein the cooling passage is from an inlet manifold at a rear portion of the combustion chamber to an outlet manifold at a front portion of the combustion chamber. It is constituted by a plurality of cooling passages extending in the axial direction of the combustion chamber, and at least one fin having two tapered surfaces is arranged in the cooling passage along the axial direction. Characteristic rocket engine.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP11696193A JP2809370B2 (en) | 1993-05-19 | 1993-05-19 | High temperature combustion engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP11696193A JP2809370B2 (en) | 1993-05-19 | 1993-05-19 | High temperature combustion engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH06330816A true JPH06330816A (en) | 1994-11-29 |
| JP2809370B2 JP2809370B2 (en) | 1998-10-08 |
Family
ID=14700032
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP11696193A Expired - Lifetime JP2809370B2 (en) | 1993-05-19 | 1993-05-19 | High temperature combustion engine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP2809370B2 (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| KR100470278B1 (en) * | 2002-07-23 | 2005-02-07 | 주식회사 로템 | rocket engine for test |
| US7373774B2 (en) | 2004-02-12 | 2008-05-20 | United Technologies Corporation | Enhanced performance torroidal coolant-collection manifold |
| WO2009140120A2 (en) * | 2008-05-15 | 2009-11-19 | Pavia, Thomas, Clayton | Systems, methods and apparatus for propulsion |
| CN112901353A (en) * | 2021-02-01 | 2021-06-04 | 中国科学院力学研究所 | System and method for starting ramjet engine through active cooling of hydrocarbon fuel and scramjet engine |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2758022C1 (en) * | 2021-02-05 | 2021-10-25 | Федеральное государственное бюджетное образовательное учреждение высшего образования Балтийский государственный технический университет "ВОЕНМЕХ" им. Д.Ф. Устинова (БГТУ "ВОЕНМЕХ") | Device for measuring the temperature of a liquid propellant rocket engine nozzle |
-
1993
- 1993-05-19 JP JP11696193A patent/JP2809370B2/en not_active Expired - Lifetime
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| KR100470278B1 (en) * | 2002-07-23 | 2005-02-07 | 주식회사 로템 | rocket engine for test |
| US7373774B2 (en) | 2004-02-12 | 2008-05-20 | United Technologies Corporation | Enhanced performance torroidal coolant-collection manifold |
| WO2009140120A2 (en) * | 2008-05-15 | 2009-11-19 | Pavia, Thomas, Clayton | Systems, methods and apparatus for propulsion |
| WO2009140120A3 (en) * | 2008-05-15 | 2010-03-04 | Pavia, Thomas, Clayton | Systems, methods and apparatus for propulsion |
| CN112901353A (en) * | 2021-02-01 | 2021-06-04 | 中国科学院力学研究所 | System and method for starting ramjet engine through active cooling of hydrocarbon fuel and scramjet engine |
Also Published As
| Publication number | Publication date |
|---|---|
| JP2809370B2 (en) | 1998-10-08 |
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