JPH0341668B2 - - Google Patents
Info
- Publication number
- JPH0341668B2 JPH0341668B2 JP58136909A JP13690983A JPH0341668B2 JP H0341668 B2 JPH0341668 B2 JP H0341668B2 JP 58136909 A JP58136909 A JP 58136909A JP 13690983 A JP13690983 A JP 13690983A JP H0341668 B2 JPH0341668 B2 JP H0341668B2
- Authority
- JP
- Japan
- Prior art keywords
- propulsion
- propellant
- nozzle
- hydrogen
- turbine
- 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
- 239000001257 hydrogen Substances 0.000 claims description 21
- 229910052739 hydrogen Inorganic materials 0.000 claims description 21
- 239000003380 propellant Substances 0.000 claims description 20
- 238000002485 combustion reaction Methods 0.000 claims description 15
- 239000007789 gas Substances 0.000 claims description 15
- 239000007788 liquid Substances 0.000 claims description 15
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 claims description 7
- 239000001301 oxygen Substances 0.000 claims description 7
- 229910052760 oxygen Inorganic materials 0.000 claims description 7
- 238000002347 injection Methods 0.000 claims description 5
- 239000007924 injection Substances 0.000 claims description 5
- 125000004435 hydrogen atom Chemical class [H]* 0.000 claims 2
- 238000001816 cooling Methods 0.000 description 12
- 150000002431 hydrogen Chemical class 0.000 description 10
- UFHFLCQGNIYNRP-UHFFFAOYSA-N Hydrogen Chemical compound [H][H] UFHFLCQGNIYNRP-UHFFFAOYSA-N 0.000 description 9
- MYMOFIZGZYHOMD-UHFFFAOYSA-N Dioxygen Chemical compound O=O MYMOFIZGZYHOMD-UHFFFAOYSA-N 0.000 description 1
- 239000003795 chemical substances by application Substances 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 238000013021 overheating Methods 0.000 description 1
- 239000000243 solution 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/46—Feeding propellants using pumps
- F02K9/48—Feeding propellants using pumps driven by a gas turbine fed by propellant combustion gases or fed by vaporized propellants or other gases
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Engine Equipment That Uses Special Cycles (AREA)
- Telescopes (AREA)
- Jet Pumps And Other Pumps (AREA)
Description
【発明の詳細な説明】
本発明は、特許請求の範囲第１項の上位概念に
記載の液体ロケツト推進装置に関する。DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a liquid rocket propulsion device according to the preamble of claim 1.
特開昭54−59516号公報（ドイツ連邦共和国特
許公開第2743983号公報）には、真空空間で駆動
するための副流構造様式の液体ロケツト推進装置
が記載されている。この液体ロケツト推進装置は
収束−発散型の前方の推進ノズル部分を有する燃
焼室と、真空推進ノズル部分と、液体推進剤、特
に酸素と水素とを送る推進剤ポンプと、推進剤ポ
ンプを駆動するのに役立ち高温駆動ガスで作用さ
れる１つ又は数個のタービンとからなり、その駆
動ガスの排気ガスが副流ノズルを経て外へ流出
し、その際タービンの駆動ガスが一方の液体推進
剤、特に水素の少量の部分量から作られ、その少
量の水素が、真空推進ノズル部分の壁内の冷却溝
を通つて流れ、そこで加熱される。 JP-A-54-59516 (German Published Patent Application No. 2743983) describes a liquid rocket propulsion device with a side flow structure for driving in a vacuum space. This liquid rocket propulsion system has a combustion chamber with a converging-divergent forward propulsion nozzle section, a vacuum propulsion nozzle section, a propellant pump for delivering liquid propellant, in particular oxygen and hydrogen, and driving a propellant pump. one or several turbines which serve and are operated with a hot motive gas, the exhaust gases of which drive gas flow out through a sidestream nozzle, with the motive gas of the turbines acting on one of the liquid propellants. , in particular from a small partial amount of hydrogen, which flows through a cooling groove in the wall of the vacuum propulsion nozzle section and is heated there.
この場合、ポンプ駆動用タービンに用いられる
駆動ガスを発生させるための汎用のジエネレータ
ーを設ける必要を無くしていて、従つて全体とし
て重量が軽減され、推進装置が安価になる。 In this case, there is no need to provide a general-purpose generator for generating the driving gas used in the pump-driving turbine, and the overall weight is therefore reduced and the propulsion device becomes less expensive.
この技術思想は、低及び中圧力比で作動するロ
ケツト推進装置には適している。高圧力比で作動
するものにはこの周知の技術思想は殆ど適用出来
ない。その理由は高エネルギーのタービン駆動ガ
スを発生するには熱の供給が充分でないからであ
る。 This concept is suitable for rocket propulsion systems operating at low and medium pressure ratios. This well-known technical concept is hardly applicable to those operating at high pressure ratios. The reason is that the heat supply is not sufficient to generate the high energy turbine driving gas.
本発明は、前述様式の液体ロケツト推進装置に
おいて、燃焼室と推進ノズルを冷却するのに役立
つ液体推進剤のために、液体循環系を作り、この
循環系がタービン駆動ガスに高い熱量を与え、推
進装置が高い圧力比で運転でき乃至は高い比出力
で作動できる様にすることを課題とするものであ
る。 The present invention provides a liquid rocket propulsion system of the type described above, in which a liquid circulation system is created for the liquid propellant that serves to cool the combustion chamber and the propulsion nozzle, and this circulation system imparts a high calorific value to the turbine driving gas; The object of the present invention is to enable a propulsion device to operate at a high pressure ratio or a high specific output.
この課題は、最初に述べた様式の推進装置にお
いて、一方の推進剤である水素の、タービン推進
ガスを発生するのに利用される少量の部分量が、
燃焼室用の噴射量として使用される大量の部分量
と共に、前方の推進ノズル部分の壁内で加熱さ
れ、次いで少量の部分量が更に加熱されるように
真空推進ノズル部分の前方端部でその壁内に導入
される様にして解決される。 This problem arises because, in the type of propulsion system mentioned at the outset, a small portion of one of the propellants, hydrogen, is used to generate the turbine propellant gas.
With the large part quantity used as injection quantity for the combustion chamber, it is heated in the wall of the forward propulsion nozzle part, and then at the forward end of the vacuum propulsion nozzle part so that the small part quantity is further heated. The solution is to introduce it into the wall.
本発明による循環系は、ポンプ駆動タービンを
駆動するための駆動ガスの温度をより高くするこ
とができ、従つて高圧力比及び高比出力の推進装
置の運転を可能にする。この目的のために与えら
れタービン側に利用される部分量を前方の推進ノ
ズル部分及び燃焼室の冷却にもいわゆる真空推進
ノズル部分の冷却にも使うことによつてこの循環
系が得られる。他方では本発明による構成は、燃
焼室と推進ノズルの冷却を所望の如くすること、
即ち前方の推進ノズル部分の後端と真空推進ノズ
ル部分の前端との間に分割平面を設定することに
より、この冷却を所望の如くすることが出来る。
即ちこの構成で、両方の推進ノズル部分間に共通
な分割平面を配置すると、非常に高温な前方の推
進ノズル部分と燃焼室及び加熱の小さい真空推進
ノズル部分の熱の条件乃至は冷却比率とが正確に
設定でき、またこの場合同時に大量の噴射推進剤
の温度並びにタービン駆動ガスの温度を容易に制
御出来るように成る。 The circulation system according to the invention allows for higher temperatures of the drive gas for driving the pump drive turbine, thus allowing operation of propulsion devices with high pressure ratios and high specific power. This circulation system is obtained in that a portion of the quantity provided for this purpose and utilized on the turbine side is also used for cooling the forward propulsion nozzle section and the combustion chamber as well as the so-called vacuum propulsion nozzle section. On the other hand, the arrangement according to the invention provides the desired cooling of the combustion chamber and the propulsion nozzle;
That is, by setting a dividing plane between the rear end of the forward propulsion nozzle section and the front end of the vacuum propulsion nozzle section, this cooling can be achieved as desired.
That is, in this configuration, by arranging a common dividing plane between both propulsion nozzle sections, the thermal conditions or cooling ratios of the front propulsion nozzle section, which is very hot, and the combustion chamber and the vacuum propulsion nozzle section, which is less heated, can be adjusted. It can be set accurately and at the same time allows easy control of the temperature of the bulk injected propellant as well as the temperature of the turbine drive gas.
次に略示した本発明による液体ロケツト推進装
置により本発明を詳細に説明する。 The invention will now be explained in more detail by means of a schematically illustrated liquid rocket propulsion device according to the invention.
液体推進剤、特に水素と酸素で作動する副流構
造様式のロケツト推進装置は、噴射ヘツド２を有
する燃焼室１と、地上又は比較的低空で駆動する
ための膨張率を有する収束−発散型の前方の推進
ノズル部分３と、空気のない空間で駆動するため
の膨張率を有する真空推進ノズル部分４と、液体
水素用の貯蔵タンク５と、液体酸素用の貯蔵タン
ク６と、水素用の推進剤ポンプ７と、酸素用の推
進剤ポンプ８と、両方のポンプ７，８を駆動する
ガスタービン９と、副流推進ノズル１０とからな
る。 A side-stream rocket propulsion system operating with liquid propellants, in particular hydrogen and oxygen, has a combustion chamber 1 with an injection head 2 and a convergent-divergent type with an expansion coefficient for operation on the ground or at relatively low altitudes. a forward propulsion nozzle section 3, a vacuum propulsion nozzle section 4 with an expansion coefficient for driving in an airless space, a storage tank 5 for liquid hydrogen, a storage tank 6 for liquid oxygen, and a propulsion nozzle section for hydrogen. It consists of an agent pump 7, a propellant pump 8 for oxygen, a gas turbine 9 that drives both pumps 7, 8, and a side stream propulsion nozzle 10.
ポンプ７から送られる全体の水素量（H2）は
導管１１を介して供給リング１２に送られ、ここ
から前方の推進ノズル部分３の壁と燃焼室１の壁
内に延在する冷却溝に供給され、冷却溝内で壁を
冷却することで水素が加熱される。全体の水素量
（H2＝H2a＋H2b）のうち少量の部分量（H2a）
が燃焼室１の前方範囲に設けた排出リング１３内
に集められ、導管１４を介して真空推進ノズル部
分４の前端の供給リング１５に供給される。 The total amount of hydrogen (H 2 ) delivered by the pump 7 is sent via a conduit 11 to a feed ring 12 and from there to a cooling groove extending in the wall of the forward propulsion nozzle part 3 and in the wall of the combustion chamber 1. The hydrogen is heated by cooling the walls in the cooling groove. A small amount (H 2 a) of the total hydrogen amount (H 2 = H 2 a + H 2 b)
is collected in a discharge ring 13 provided in the front region of the combustion chamber 1 and is fed via a conduit 14 to a feed ring 15 at the front end of the vacuum propulsion nozzle part 4 .
水素（H2）の大量の部分量（H2b）は導管１
６を介して噴射ヘツド２内に流入し、そこで導管
１７を介して流入する酸素（O2）と混合され、
燃焼室１内に噴射される。 A large partial quantity (H 2 b) of hydrogen (H 2 ) is
6 into the injection head 2 where it is mixed with oxygen (O 2 ) entering via conduit 17;
is injected into the combustion chamber 1.
供給リング１５内に流入され既に加熱された少
量の水素量（H2a）はそこで真空推進ノズル部分
４の壁を形成する管束に分配され、真空推進ノズ
ル部分４の後端で排出リング１８に集められる。
真空推進ノズル部分４内では水素（H2）の少量
部分量（H2a）が付加的に加熱される。水素
（H2）のタービン側で利用される少量の部分量の
熱いガスは導管１９を介してタービン９に達し、
タービンを駆動して副流推進ノズル１０を経て自
由空間に流出する。 The small amount of hydrogen (H 2 a) which has entered the supply ring 15 and has already been heated is distributed there to the tube bundle forming the wall of the vacuum-propelled nozzle section 4 and is transferred to the discharge ring 18 at the rear end of the vacuum-propelled nozzle section 4 . Can be collected.
In the vacuum propulsion nozzle section 4 a small portion (H 2 a) of hydrogen (H 2 ) is additionally heated. A small partial quantity of hot gas utilized on the turbine side of hydrogen (H 2 ) reaches the turbine 9 via a conduit 19;
It drives the turbine and flows out into free space through the side stream propulsion nozzle 10.
本発明は、この実施例でしめした様に、冷却の
ため使用される全体の推進剤量、即ち水素量
（H2＋H2b）が燃焼室を含めた前方の推進ノズル
部分の壁内で加熱され、そしてタービンにとつて
必要な少量の部分量（H2a）が引き続いて更に真
空推進ノズル部分の壁内で加熱されるので、ター
ビン駆動ガスに高い熱量が与えられ、ロケツト推
進装置は高い圧力比乃至は高い比出力で運転でき
るようになる。 As shown in this embodiment, the present invention is such that the total amount of propellant used for cooling, that is, the amount of hydrogen (H 2 + H 2 b) is within the wall of the forward propulsion nozzle portion including the combustion chamber. The small partial quantity (H 2 a) that is heated and required for the turbine is subsequently further heated in the walls of the vacuum propulsion nozzle section, so that a high calorific value is imparted to the turbine drive gas and the rocket propulsion system It becomes possible to operate at a high pressure ratio or high specific power.
また最初に述べた既に公知の従来技術である特
開昭54−59516号公報においては、冷却推進剤が
最初から前方の推進ノズル部分及び燃焼室と、後
方の推進ノズル部分とに分割されるため、即ち２
つの部分流に分かれて供給されるために高い比出
力が得られないが、本発明によりこの点が改良さ
れるという長所を得ることが出来る。更にロケツ
ト推進装置の高い比出力は、推進装置の冷却が最
適となる時にのみ確実に得られるが、これも全体
の冷却推進剤（H2a＋H2b）が先ず最大の熱負荷
のかかる推進剤部分、即ち前方の推進ノズル部分
及び燃焼室を冷却する様にして最適の冷却が得ら
れる長所も生ずる。従つて過熱の危険、即ち構成
部材が燃えるという危険は本発明による構成で完
全に避けることが出来る。尚、熱的にそれ程負荷
を受けない真空推進ノズル部分を冷却するために
は、予め加熱した推進剤の少量の部分量H2aで充
分である。 Furthermore, in the already well-known prior art disclosed in Japanese Patent Application Laid-Open No. 54-59516, the cooled propellant is divided from the beginning into the forward propulsion nozzle portion and combustion chamber, and the rear propulsion nozzle portion. , i.e. 2
Although a high specific power cannot be obtained because the supply is divided into two partial streams, the present invention has the advantage of improving this point. Furthermore, the high specific power of a rocket propulsion system can only be achieved reliably when the cooling of the propulsion system is optimal, which also means that the total cooled propellant (H 2 a + H 2 b) is first of all the propellant with the greatest thermal load. The advantage also arises that optimum cooling is achieved in such a way that parts, ie the forward propulsion nozzle part and the combustion chamber, are cooled. The risk of overheating, ie the risk of burning out the components, can therefore be completely avoided with the design according to the invention. It should be noted that a small partial quantity H 2 a of preheated propellant is sufficient for cooling the vacuum propulsion nozzle parts which are not thermally heavily loaded.
図面は本発明の一実施例を説明する略図であ
る。
図中参照番号、１……燃焼室、２……燃料噴射
ヘツド、３……前方の推進ノズル部分、４……真
空推進ノズル部分、５……水素タンク、６……酸
素タンク、７……水素ポンプ、８……酸素ポン
プ、９……タービン、１０……副流推進ノズル、
H2……水素全体量、H2a……少量の水素部分量、
H2b……大量の水素部分量。
The drawing is a schematic diagram illustrating an embodiment of the invention. Reference numbers in the figure: 1... Combustion chamber, 2... Fuel injection head, 3... Front propulsion nozzle part, 4... Vacuum propulsion nozzle part, 5... Hydrogen tank, 6... Oxygen tank, 7... Hydrogen pump, 8...Oxygen pump, 9...Turbine, 10...Sidestream propulsion nozzle,
H 2 ... total amount of hydrogen, H 2 a ... small amount of hydrogen,
H 2 b...a large amount of hydrogen.
Claims (1)
式の液体ロケツト推進装置にして、その推進装置
が主として収束−発散型の前方の推進ノズル部分
を有する燃焼室と、その推進ノズル部分に続く真
空推進ノズル部分と、液体の推進剤、特に酸素と
水素とを送るための推進剤ポンプと、推進剤ポン
プを駆動するのに役立つ１つ又は数個のタービン
とからなり、それらタービンが高温のタービン駆
動ガスの作用を受け、その駆動ガスは推進ノズル
壁の内部で熱の供給を受けて一方の推進剤、特に
水素の部分量から作られ、次いでタービンを通過
し、タービン排気ガスとして副流ノズルを経て自
由空間へ出る様な、液体ロケツト推進装置におい
て、一方の推進剤である水素の、タービン推進ガ
スを発生するのに利用される少量の部分量
（H2a）が、燃焼室１用の噴射量として使用され
る大量の部分量（H2b）と共に、前方の推進ノズ
ル部分３の壁内で加熱され、次いで少量の部分量
（H2a）が更に加熱されるように真空推進ノズル
部分４の前方端部でその壁内に導入されることを
特徴とする液体ロケツト推進装置。1. A liquid rocket propulsion device with a side-flow structure for operation in an airless space, the propulsion device comprising a combustion chamber having a primarily convergent-divergent forward propulsion nozzle portion, and a vacuum following the propulsion nozzle portion. It consists of a propellant nozzle section, a propellant pump for delivering liquid propellant, in particular oxygen and hydrogen, and one or several turbines serving to drive the propellant pump, which turbines are connected to a hot turbine. Under the action of the driving gas, which is produced from a partial quantity of one propellant, in particular hydrogen, with a supply of heat inside the propulsion nozzle wall, it then passes through the turbine and is discharged as turbine exhaust gas to the sidestream nozzle. In a liquid rocket propulsion device that exits into free space through The vacuum propulsion is heated in the wall of the forward propulsion nozzle section 3, with a large part quantity (H 2 b) used as the injection quantity, and then a small part quantity (H 2 a) is heated further. Liquid rocket propulsion device, characterized in that it is introduced at the forward end of the nozzle part 4 into its wall.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| DE19823228162 DE3228162A1 (en) | 1982-07-28 | 1982-07-28 | Liquid-fuelled rocket motor of the subsidiary-flow type, for operation in space where there is no air |
| DE3228162.5 | 1982-07-28 |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS5941645A JPS5941645A (en) | 1984-03-07 |
| JPH0341668B2 true JPH0341668B2 (en) | 1991-06-24 |
Family
ID=6169517
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP13690983A Granted JPS5941645A (en) | 1982-07-28 | 1983-07-28 | Sub-current structure type liquid rocket drive for driving in vacuum space |
Country Status (3)
| Country | Link |
|---|---|
| JP (1) | JPS5941645A (en) |
| DE (1) | DE3228162A1 (en) |
| FR (1) | FR2531141B1 (en) |
Families Citing this family (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB8610849D0 (en) * | 1986-05-02 | 1986-08-20 | Marconi Co Ltd | Gas thruster |
| JPS62261652A (en) * | 1986-05-07 | 1987-11-13 | Natl Space Dev Agency Japan<Nasda> | Liquid rocket engine |
| US4771600A (en) * | 1986-10-20 | 1988-09-20 | United Technologies Corporation | Tripropellant rocket engine |
| US5267437A (en) * | 1991-05-23 | 1993-12-07 | United Technologies Corporation | Dual mode rocket engine |
| RU2450153C1 (en) * | 2011-02-07 | 2012-05-10 | Александр Фролович Ефимочкин | Liquid propellant rocket engine |
| FR2981127B1 (en) | 2011-10-11 | 2013-11-29 | Snecma | REACTION PROPULSION DEVICE AND FEEDING METHOD |
| FR3012848B1 (en) * | 2013-11-06 | 2015-11-27 | Snecma | PROPELLANT ASSEMBLY AND PROCESS FOR SUPPLYING ERGOLS |
| RU2554126C1 (en) * | 2013-12-18 | 2015-06-27 | Федеральное государственное унитарное предприятие "Государственный космический научно-производственный центр имени М.В. Хруничева" | Combined engine unit of rocket pod |
| RU2612512C1 (en) * | 2016-03-29 | 2017-03-09 | Владислав Юрьевич Климов | Liquid propellant rocket engine |
| CN111409877A (en) * | 2020-03-13 | 2020-07-14 | 上海空间推进研究所 | Air suction and separation device for aircraft hatch filling pipeline |
| RU2760956C1 (en) * | 2020-11-10 | 2021-12-01 | Акционерное общество "КБхиммаш им. А.М. Исаева" | Liquid rocket engine with an electric pump supply system |
| CN114136635B (en) * | 2021-12-06 | 2022-08-23 | 北京航空航天大学 | Large-flow quick-response solid-liquid rocket engine ground conveying system |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS6131296A (en) * | 1984-07-24 | 1986-02-13 | 京セラミタ株式会社 | Custody facility for recording materials such as document |
Family Cites Families (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3077073A (en) * | 1957-10-29 | 1963-02-12 | United Aircraft Corp | Rocket engine having fuel driven propellant pumps |
| US3267664A (en) * | 1963-03-19 | 1966-08-23 | North American Aviation Inc | Method of and device for cooling |
| US4171615A (en) * | 1966-04-21 | 1979-10-23 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Supercharged topping rocket propellant feed system |
| DE2743983C2 (en) * | 1977-09-30 | 1982-11-11 | Messerschmitt-Bölkow-Blohm GmbH, 8000 München | By-pass liquid rocket engine for operation in a vacuum |
-
1982
- 1982-07-28 DE DE19823228162 patent/DE3228162A1/en active Granted
-
1983
- 1983-07-21 FR FR8312095A patent/FR2531141B1/en not_active Expired
- 1983-07-28 JP JP13690983A patent/JPS5941645A/en active Granted
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS6131296A (en) * | 1984-07-24 | 1986-02-13 | 京セラミタ株式会社 | Custody facility for recording materials such as document |
Also Published As
| Publication number | Publication date |
|---|---|
| JPS5941645A (en) | 1984-03-07 |
| DE3228162C2 (en) | 1987-06-19 |
| FR2531141B1 (en) | 1987-03-20 |
| DE3228162A1 (en) | 1984-02-09 |
| FR2531141A1 (en) | 1984-02-03 |
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