EP0401107A1 - Brennkammer für Staustrahltriebwerk - Google Patents

Brennkammer für Staustrahltriebwerk Download PDF

Info

Publication number
EP0401107A1
EP0401107A1 EP90401425A EP90401425A EP0401107A1 EP 0401107 A1 EP0401107 A1 EP 0401107A1 EP 90401425 A EP90401425 A EP 90401425A EP 90401425 A EP90401425 A EP 90401425A EP 0401107 A1 EP0401107 A1 EP 0401107A1
Authority
EP
European Patent Office
Prior art keywords
chamber
flow
injection
injection device
wall
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
Application number
EP90401425A
Other languages
English (en)
French (fr)
Other versions
EP0401107B1 (de
Inventor
Philippe H. Ramette
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Societe Europeenne de Propulsion SEP SA
Original Assignee
Societe Europeenne de Propulsion SEP SA
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Societe Europeenne de Propulsion SEP SA filed Critical Societe Europeenne de Propulsion SEP SA
Publication of EP0401107A1 publication Critical patent/EP0401107A1/de
Application granted granted Critical
Publication of EP0401107B1 publication Critical patent/EP0401107B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/346—Feeding into different combustion zones for staged combustion

Definitions

  • the present invention relates to a supersonic combustion ramjet.
  • Supersonic combustion ramjet engines are currently being studied for the propulsion of hypersonic vehicles, for example recoverable space planes with horizontal takeoff.
  • the propulsion phase by supersonic combustion ramjet makes it possible to accelerate the vehicle by the speed - approximately Mach 6 - reached at the end of the propulsion phase by subsonic combustion ramjet, up to a speed of approximately Mach 15 to Mach 25.
  • the air circulates at a speed which is always supersonic in the middle of the air stream, where the wall effects are hardly felt, and the fuel, generally hydrogen gas. is introduced through the wall of the chamber.
  • the injection of the hydrogen gas flow is generally carried out by holes or slots formed in the wall of the chamber. It is difficult to ensure a satisfactory mixture between hydrogen and air, and therefore to obtain good energy efficiency, without aerodynamic flow losses due to interactions or shocks between the air flow and the flow.
  • injected hydrogen In fact, an injection of hydrogen through holes directed towards the axis of the combustion chamber necessarily produces shocks between the gas flows.
  • the hydrogen is injected tangentially to the wall of the chamber, it tends to remain confined against it under the effect of the air flowing at high speed in the chamber, and combustion occurs. incomplete because of the short air residence time in the room.
  • the present invention aims to provide a ramjet chamber with supersonic combustion into which a flow of gaseous fuel can be introduced without creating damaging shocks with the air flowing in the chamber, at speed. supersonic, while obtaining satisfactory energy efficiency.
  • a ramjet comprising a combustion chamber intended to be traversed longitudinally by an air flow at supersonic speed, and a first injection device for injecting into the chamber a flow of gaseous fuel with a speed entry into the chamber having a low amplitude transverse component
  • ramjet in which a second injection device is located downstream of the first, in the direction of the flow of air at supersonic speed, for injecting into the chamber a flow of gaseous oxidizer which contributes to detaching from the wall of the chamber the flow of gaseous fuel injected by the first injection device.
  • the first injection device preferably comprises a first wall part of the combustion chamber, for example a ring-shaped part, which is made of a material permeable to the flow of gaseous fuel to be injected into the chamber and which has a surface constituting a part of the interior surface of the chamber and an opposite surface in communication with a source of the gaseous fuel to be injected, so that the injection of the gaseous fuel flow is carried out by transpiration through the porosity of the porous material constituting of the first injection device.
  • a first wall part of the combustion chamber for example a ring-shaped part, which is made of a material permeable to the flow of gaseous fuel to be injected into the chamber and which has a surface constituting a part of the interior surface of the chamber and an opposite surface in communication with a source of the gaseous fuel to be injected, so that the injection of the gaseous fuel flow is carried out by transpiration through the porosity of the porous material constituting of the first injection device.
  • the second injection device can be produced in the same way.
  • porous material through which the gas flow transpires is an injection means which is perfectly suitable for injecting the gas flow into the chamber with an input speed having a radial component of low amplitude.
  • the porous material is advantageously a porous composite material with a ceramic or carbon matrix.
  • a material is particularly suitable for producing a device for injecting a gas flow into a ramjet combustion chamber.
  • thermostructural that is to say a mechanical behavior at high temperature which makes it possible to produce an injection device constituting a structural element of the chamber.
  • the porosity of this material can be controlled by acting on the volume ratio of fibers constituting its fibrous reinforcing texture and / or on the degree of densification by the material constituting the matrix.
  • a material of type C / SiC (reinforcement of carbon fibers and matrix of silicon carbide), or of type SiC / SiC (reinforcement of fibers essentially of silicon carbide and matrix of silicon carbide), or of type C / C protected (carbon fiber reinforcement, carbon matrix and anti-oxidation protection), may be suitable.
  • the wall of the chamber at least in its parts adjacent to the injection devices, is also made of non-porous composite material with ceramic or carbon matrix.
  • the connection between the injection devices and the other parts of the wall of the combustion chamber can then advantageously be carried out by co-densification of the wall parts forming injection devices and of the other wall parts assembled in an incomplete state. densified. This co-densification is preferably carried out by chemical vapor deposition.
  • Injection methods other than by transpiration through a porous material may be used to inject the flow of gaseous fuel or the flow of gaseous oxidant.
  • the fuel flow must be injected with a low speed radial component so as not to cause violent interactions or shocks with the air flow at supersonic speed; it is preferably the same for the injection of the gaseous oxidant flow.
  • Injectors or injection orifices opening into the chamber substantially tangentially to the wall thereof may be provided.
  • the chamber 10 is of cylindrical shape with circular section and comprises, in the direction of air flow at supersonic speed (arrow A), an upstream sealed section 12, a first injection ring 20 for the injection of a flow of gaseous fuel, a central sealed section 14, a second injection ring 30 for the injection of a flow of gaseous oxidizer and a downstream sealed section 16.
  • the interior surfaces of the sections 12, 14, 16 and injection rings 20, 30 define the continuous cylindrical internal wall of the ramjet chamber.
  • the outer surface of the ring 20 defines a chamber 22 for injecting gaseous fuel which communicates with a fuel source (not shown).
  • the fuel is for example hydrogen which is injected in the gaseous state, the pressure prevailing in the injection chamber 22 being greater than that prevailing in the combustion chamber of the ramjet.
  • the ring 20 is made in a single piece of porous refractory composite material.
  • the porosity of the material constituting the ring 20 gives the latter the permeability necessary to allow the injection of the gaseous flow of hydrogen by transpiration through the injection ring.
  • the hydrogen gas flow thus enters the chamber with a low radial velocity component.
  • the flow of hydrogen injected into the combustion chamber is defined by the porosity of the injection ring, the length of the latter, and the pressure difference between the outer and inner surfaces of the ring.
  • the constituent material of the ring 20 is a composite material consisting of a refractory fibrous reinforcement partially densified by a ceramic material, or of a fibrous carbon reinforcement partially densified by a carbon matrix and protected against oxidation.
  • a refractory fibrous reinforcement partially densified by a ceramic material
  • a fibrous carbon reinforcement partially densified by a carbon matrix and protected against oxidation.
  • the preform is made of carbon fibers or ceramic fibers, for example fibers essentially of silicon carbide.
  • the fiber preform is produced by winding on a mandrel of a strip of fabric until the desired thickness is obtained.
  • the superimposed layers of fabric can be linked together by needling or implantation of threads.
  • the preform is densified by gas or by liquid.
  • a matrix is produced by chemical vapor infiltration of ceramic material, for example silicon carbide, or carbon (for a protected C / C type material).
  • the preform is impregnated with a precursor of the matrix material, which is then obtained by heat treatment.
  • an injection ring made of ceramic material C / SiC can be produced by manufacturing a carbon fiber preform having a fiber volume ratio of approximately 35% and densifying it by chemical vapor infiltration of silicon carbide until reaching a residual porosity of about 40
  • the ring 30 delimits by its outer surface a chamber 32 for injecting gaseous oxidant.
  • This can consist of air taken from the surrounding medium or oxygen from a source (not shown).
  • the ring 30 is made in a single piece of porous composite material either with a ceramic matrix, for example in C / SiC material, or of C / C type protected against oxidation, in the same way as the ring 20.
  • the porosity of the material of the ring 30 gives the latter the permeability necessary to allow the injection of a flow of gaseous oxidizer by transpiration through the ring 30, the pressure in the injection chamber 32 being greater than that prevailing in the chamber ramjet.
  • the flow of gaseous oxidizer into the chamber is therefore also carried out with a low radial velocity component.
  • the sections 12, 14, 16 of the ramjet chamber are preferably also made of a composite material with a ceramic or carbon matrix.
  • a material having a reinforcement and a matrix of the same type as those of the injection rings 20 and 30 will be chosen.
  • the sections 12, 14 and 16 are sealed, the seal being obtained by densification sufficiently advanced to fill the porosity of the fibrous reinforcement until the material is impermeable.
  • connection between the sections 12, 14, 16 of the wall of the chamber 10 and the injection rings 20, 30 is produced by co-densification.
  • the sections 12, 14, 16 and the rings 20, 30 are produced separately while being incompletely densified with respect to the desired degree of final densification.
  • the elements are then assembled end to end and placed in an infiltration oven to undergo a final co-densification by chemical vapor infiltration.
  • the continuity of the matrix material at the interfaces between the sections 12, 14, 16 and the rings 20, 30 ensures the connection between these elements.
  • This final co-densification is continued until the desired degree of porosity is obtained for the injection rings 20 and 30, the sections 12, 14, 16 having previously been sufficiently pre-densified to finally obtain the desired seal.
  • the gas flow 34 of oxidant transpiring through the injection ring 30 forces the gas flow 24 of fuel to move away from the wall of the chamber 10 despite the supersonic air flow having tendency to press it against this wall.
  • a satisfactory mixture is thus obtained between the combustible gas and the oxidizer constituted by the supersonic air and the flow 34.
  • Complete combustion of the combustible gas can thus be carried out in a very short time, without creating violent interactions between the current d supersonic air and gas flows transpiring through the injection rings. This results in an increase in performance of the ramjet chamber, therefore better thrust and specific impulse of the propulsion system.
  • porous materials for example porous metallic structures, can be used in the case of a metallic chamber.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Fuel-Injection Apparatus (AREA)
  • Ceramic Products (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
EP19900401425 1989-05-29 1990-05-29 Brennkammer für Staustrahltriebwerk Expired - Lifetime EP0401107B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8907019 1989-05-29
FR8907019A FR2647533B1 (fr) 1989-05-29 1989-05-29 Chambre de statoreacteur a combustion supersonique

Publications (2)

Publication Number Publication Date
EP0401107A1 true EP0401107A1 (de) 1990-12-05
EP0401107B1 EP0401107B1 (de) 1993-07-21

Family

ID=9382105

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19900401425 Expired - Lifetime EP0401107B1 (de) 1989-05-29 1990-05-29 Brennkammer für Staustrahltriebwerk

Country Status (4)

Country Link
EP (1) EP0401107B1 (de)
JP (1) JPH0396645A (de)
DE (1) DE69002281T2 (de)
FR (1) FR2647533B1 (de)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1999004156A1 (de) * 1997-07-17 1999-01-28 Deutsches Zentrum für Luft- und Raumfahrt e.V. Brennkammer und verfahren zur herstellung einer brennkammer
FR2836699A1 (fr) * 2002-03-04 2003-09-05 Eads Launch Vehicles Moteur de fusee
FR2836698A1 (fr) * 2002-03-04 2003-09-05 Eads Launch Vehicles Chambre de combustion pour statoreacteur et statoreacteur pourvu d'une telle chambre de combustion
CN103343983A (zh) * 2013-07-31 2013-10-09 哈尔滨工业大学 基于强磁场稳定电弧的超声速稳定燃烧方法
GB2518211A (en) * 2013-09-13 2015-03-18 Carolyn Billie Knight Evaporative wick/membrane rocket motor
CN108317541A (zh) * 2018-02-26 2018-07-24 中国科学院力学研究所 一种冲压发动机

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP4522558B2 (ja) * 2000-08-11 2010-08-11 実 屋我 スクラムジェットエンジン用燃料混合促進方法並びに装置
CN113530709B (zh) * 2021-09-16 2021-12-14 西安空天引擎科技有限公司 一种双模态过氧化氢燃气发生器

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2658332A (en) * 1951-03-21 1953-11-10 Carborundum Co Fluid cooled, refractory, ceramic lined rocket structure
US3114961A (en) * 1959-03-20 1963-12-24 Power Jets Res & Dev Ltd Treatment of porous bodies
GB1046909A (en) * 1963-08-26 1966-10-26 Gur Charan Saini Rocket thrust chambers
DE1278319B (de) * 1963-11-28 1969-04-17 Bbc Brown Boveri & Cie Verfahren zum Schuetzen von durch heisse Medien ueberstrichenen Oberflaechenteilen eines hitzebestaendigen Koerpers
FR2158572A1 (de) * 1971-11-05 1973-06-15 Penny Robert
US3864907A (en) * 1973-11-05 1975-02-11 Us Air Force Step cylinder combustor design
GB2053873A (en) * 1979-07-19 1981-02-11 Europ Propulsion High temperature thermal insulation material and method for making same
GB2089434A (en) * 1980-12-09 1982-06-23 Rolls Royce Composite Ducts for Jet Pipes
GB2196393A (en) * 1986-10-14 1988-04-27 Gen Electric Propulsion apparatus and method

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE253189C (de) *

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2658332A (en) * 1951-03-21 1953-11-10 Carborundum Co Fluid cooled, refractory, ceramic lined rocket structure
US3114961A (en) * 1959-03-20 1963-12-24 Power Jets Res & Dev Ltd Treatment of porous bodies
GB1046909A (en) * 1963-08-26 1966-10-26 Gur Charan Saini Rocket thrust chambers
DE1278319B (de) * 1963-11-28 1969-04-17 Bbc Brown Boveri & Cie Verfahren zum Schuetzen von durch heisse Medien ueberstrichenen Oberflaechenteilen eines hitzebestaendigen Koerpers
FR2158572A1 (de) * 1971-11-05 1973-06-15 Penny Robert
US3864907A (en) * 1973-11-05 1975-02-11 Us Air Force Step cylinder combustor design
GB2053873A (en) * 1979-07-19 1981-02-11 Europ Propulsion High temperature thermal insulation material and method for making same
GB2089434A (en) * 1980-12-09 1982-06-23 Rolls Royce Composite Ducts for Jet Pipes
GB2196393A (en) * 1986-10-14 1988-04-27 Gen Electric Propulsion apparatus and method

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1999004156A1 (de) * 1997-07-17 1999-01-28 Deutsches Zentrum für Luft- und Raumfahrt e.V. Brennkammer und verfahren zur herstellung einer brennkammer
US6151887A (en) * 1997-07-17 2000-11-28 Deutsches Zentrum Fuer Luft- Und Raumfahrt E.V. Combustion chamber for rocket engine
FR2836699A1 (fr) * 2002-03-04 2003-09-05 Eads Launch Vehicles Moteur de fusee
FR2836698A1 (fr) * 2002-03-04 2003-09-05 Eads Launch Vehicles Chambre de combustion pour statoreacteur et statoreacteur pourvu d'une telle chambre de combustion
EP1342905A1 (de) * 2002-03-04 2003-09-10 Eads Launch Vehicles Raketenmotor
EP1342904A1 (de) * 2002-03-04 2003-09-10 Eads Launch Vehicles Brennkammer für ein Staustrahltriebwerk und Staustrahltriebwerk mit einer solchen Brennkammer
WO2003074858A2 (fr) * 2002-03-04 2003-09-12 Eads Space Transportation Sa Chambre de combustion pour statoréacteur et statoréacteur pourvu d'un telle chambre de combustion
WO2003074859A1 (fr) * 2002-03-04 2003-09-12 Eads Space Transportation Sa Moteur de fusee
WO2003074858A3 (fr) * 2002-03-04 2004-04-01 Eads Space Transportation Sa Chambre de combustion pour statoréacteur et statoréacteur pourvu d'un telle chambre de combustion
US6915627B2 (en) 2002-03-04 2005-07-12 Eads Space Transportation Sa Rocket engine
US7000398B2 (en) 2002-03-04 2006-02-21 Eads Space Transportation Sa Ramjet engine combustion chamber and ramjet engine equipped with same
CN103343983A (zh) * 2013-07-31 2013-10-09 哈尔滨工业大学 基于强磁场稳定电弧的超声速稳定燃烧方法
CN103343983B (zh) * 2013-07-31 2014-12-24 哈尔滨工业大学 基于强磁场稳定电弧的超声速稳定燃烧方法
GB2518211A (en) * 2013-09-13 2015-03-18 Carolyn Billie Knight Evaporative wick/membrane rocket motor
GB2518211B (en) * 2013-09-13 2015-11-18 Carolyn Billie Knight Rocket motor with combustion chamber having porous membrane
CN108317541A (zh) * 2018-02-26 2018-07-24 中国科学院力学研究所 一种冲压发动机
CN108317541B (zh) * 2018-02-26 2020-07-07 中国科学院力学研究所 一种冲压发动机

Also Published As

Publication number Publication date
FR2647533B1 (fr) 1993-03-19
DE69002281D1 (de) 1993-08-26
DE69002281T2 (de) 1994-01-27
FR2647533A1 (fr) 1990-11-30
EP0401107B1 (de) 1993-07-21
JPH0396645A (ja) 1991-04-22

Similar Documents

Publication Publication Date Title
EP0604279B1 (de) Injektor mit einer porösen Wand für eine Raketenbrennkammer
EP0817762B1 (de) Durch selbstheilende matrix oxidationsgeschütztes verbundmaterial und verfahren zu seiner herstellung
EP1045971B1 (de) Wärmetauscher aus verbundwerkstoff und dessen herstellungsverfahren
EP2132426A1 (de) Cmc-mischer mit struktureller aussenverschalung
EP0401107B1 (de) Brennkammer für Staustrahltriebwerk
EP1342905B1 (de) Raketenmotor
FR2944062A1 (fr) Injecteur d'ergols
FR2791589A1 (fr) Procede de fabrication d'une tuyere de moteur de fusee a refroidissement et tuyere obtenue
EP1342904B1 (de) Brennkammer für ein Staustrahltriebwerk und Staustrahltriebwerk mit einer solchen Brennkammer
CA2429393A1 (fr) Perfectionnement aux procedes de densification par calefaction d'une structure poreuse
Patterson et al. Advanced HfC-TaC oxidation resistant composite rocket thruster
EP0421865A1 (de) Raketenbrennkammer
EP0401106B1 (de) Reaktionskammer und Verfahren zu ihrer Herstellung
WO1993013636A1 (fr) Procede de formation d'un passage etanche dans une piece en materiau composite refractaire, et application a la realisation d'une structure en materiau composite refractaire refroidie par circulation de fluide
FR2782378A1 (fr) Piece de structure comportant une partie en materiau composite thermostructural refroidie par circulation de fluide
WO2019220057A1 (fr) Procede de fabrication d'une piece cmc
FR3113496A1 (fr) Procédé de dépôt d’un revêtement sur un fil sous champ micro-ondes
WO1995007768A1 (fr) Procede pour la realisation de materiaux ou revetements composites et installation pour sa mise en ×uvre
FR3081156A1 (fr) Procede de fabrication d'une piece cmc revetue
WO2024084152A1 (fr) Preforme fibreuse et son procede de fabrication pour realiser une piece en materiau composite a matrice ceramique
EP4192690A1 (de) Verfahren zur generativen fertigung zur herstellung einer struktur
FR3084445A1 (fr) Fabrication d'une chambre de combustion en materiau composite
FR3027959B1 (fr) Protection anti-feu d'une piece en materiau composite d'une turbine a gaz
FR2645071A1 (fr) Procede pour la realisation de trous dans une piece en materiau composite
FR3136463A1 (fr) Procédé de traitement d’au moins une fibre en céramique ou en carbone

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): DE GB

17P Request for examination filed

Effective date: 19901217

17Q First examination report despatched

Effective date: 19920109

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE GB

REF Corresponds to:

Ref document number: 69002281

Country of ref document: DE

Date of ref document: 19930826

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)

Effective date: 19931018

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 20000512

Year of fee payment: 11

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20000523

Year of fee payment: 11

REG Reference to a national code

Ref country code: GB

Ref legal event code: 732E

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20010529

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20010529

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20020301