US3719046A - Rocket engine cooling system - Google Patents

Rocket engine cooling system Download PDF

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US3719046A
US3719046A US3719046DA US3719046A US 3719046 A US3719046 A US 3719046A US 3719046D A US3719046D A US 3719046DA US 3719046 A US3719046 A US 3719046A
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Prior art keywords
coolant
chamber
primary
gases
propulsive
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Expired - Lifetime
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G Sutherland
D Emmons
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ROCKET RESEARCH CORP
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ROCKET RESEARCH CORP
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60Constructional parts; Details
    • F02K9/62Combustion or thrust chambers
    • F02K9/64Combustion or thrust chambers having cooling arrangements
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60Constructional parts; Details
    • F02K9/68Decomposition chambers
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • F05D2260/208Heat transfer, e.g. cooling using heat pipes

Abstract

An oxidizer, such as nitrogen tetroxide or fluorine, is introduced into a reaction chamber in the path of the decomposition products of hydrazine and reacts therewith to form high temperature, high thrust propulsive gases. A heat pipe surrounds the reaction chamber and includes a wick saturated with a volatile liquid, such as liquid lithium, which liquid is vaporized thereby removing heat from the chamber wall. The vaporized fluid is directed through a heat exchanger and is therein condensed back into a liquid state.

Description

United States Patent 1 1 1111 3,719,046

Sutherland et al. 51 March 6, 1973 54] ROCKET ENGINE COOLING SYSTEM 3,149,460 9 1964 ROCCa ..60/260 3,232,048 2 1966 Stockel .60/261 [75] Invemms Gmrge Sulherland, Meme 3,493,177 2/1970 Bromberg .:.60/267 Island; Donald L. Emmons, ls-

Saquah both of wash Primary ExaminerDouglas Hart [73] Assignee: Rocket Research Corporation, yfi Barnard, Uhlir & Hughes Redmond, Wash.

[22] Filed: July 2, 1970 [21] Appl. No.: 52,020

[5 7 ABSTRACT An oxidizer, such as nitrogen tetroxide or fluorine, is introduced into a reaction chamber in the path of the decomposition products of hydrazine and reacts [52] us. Cl ..60/206, 60/39.46, 60/39.5l R, therewith to form g temperature, g thrust 60/39.66, 60/267, 60/207, 60/224 propulsive gases. A heat pipe surrounds the reaction [51] Int. Cl. ..F02k 9/02, F02k 1 1/02 chamber and includes a wick saturated with a volatile [58] Field of Search ..60/260, 267, 204, 206, 207, liquid, such as liquid lithium, which liquid is vaporized 60/224, 39.46, 39.51 R, 39.66, 261, 37, thereby removing heat from the chamber wall. The 39.12; 165/ 105 vaporized fluid is directed through a heat exchanger and is therein condensed back into a liquid state. [5 6] References Cited 13 Claims, 5 Drawing Figures UNITED STATES PATENTS 3,024,606 3/1962 Adams ..60/267 ot/ n 5, w

. a /4 Fl/EL SOUEZ'E I g 2 a" Z e 20 I 24 I I I I b 1 42 i I V 40 OXIDIZE /7 SOURCE PATENTEDHAR 61m SHEET 1 BF 3 muwsom mNxo PATE EW 3.719.046

sum 2 OF 3 AWTOP/VEYS PATENTEUHAR 6 1975 3 ,719,046

SHEET 30F 3 ROCKET ENGINE COOLING SYSTEM BACKGROUND OF THE INVENTION first stage are reacted with an oxidizer in the second stage, and to the use of a liquid propellant or some of the low temperature products of the first stage in a heat exchanger to condense the coolant within the heat pipe.

2. Description of the Prior Art In a heat pipe cooling system a liquid coolant is carried in a wick and is directed into a high temperature zone in which the coolant is vaporized thereby removing heat from the high temperature source. The vaporizer coolant then flows to a heat exchanger where it is condensed back into a liquid. The liquid returns to the high temperature zone by moving through the wick under the influence of capillary action. Reference is made to the detailed explanation of the heat pipe principle presented in the May, 1968 issue of Scientific America, commencing at page 38.

SUMMARY OF THE INVENTION This invention advantageously combines heat pipe cooling principles with a two-stage hydrazone engine concept to provide a compact, high performance bipropellant rocket engine capable of long duration operation.

According to the invention the hydrazone is decomposed in a first stage reaction chamber with the said products of decomposition being used in a heat exchanger to condense a vaporizer coolant in a heat pipe. The latter is employed to cool a second stage reaction chamber in which the decomposed hydrazine is reacted with an oxidizer producing high temperature propulsive gases.

In another form of the invention a liquid propellant is used to condense the vaporized heat pipe coolant.

BRIEF DESCRIPTION OF THE DRAWING FIG. I is a longitudinal sectional view of a two-stage rocket engine embodying certain heat pipe cooling principles of this invention;

FIG. 2 is a fragmentary sectional view taken through the heat exchanger substantially along line 2-2 of FIG.

FIG. 3 is a fragmentary sectional view transversely of the engine, taken substantially along line 33 of FIG.

FIG. 4 is a longitudinal sectional view of a modified form of engine having a film cooled chamber and a heat pipe cooled nozzle; and

FIG. 5 is a longitudinal sectional view of another modified form of engine using both film and heat pipe cooling for both the reaction chamber and the nozzle.

DESCRIPTION OF THE PREFERRED EMBODIMENTS The rocket engine of FIG. 1 comprises a reaction chamber wall 12 of stainless steel or the like terminating at its downstream end in a nozzle section 14. The

chamber is divided internally into a first stage decomposition chamber 16 and a second stage reaction chamber 17, axially separated by a perforated wall 18. A fuel injector 20 introduces liquid hydrazine into the first stage reaction chamber 16. The hydrazine reacts with a catalyst, such as Shell 405, identified by the reference character 22, to form products of decomposition at a temperature of about 1,600F. For higher thrust, these products of decomposition are reacted with an oxidizer (e.g. nitrogen tetroxide, N 0 fluorine; oxygen; flox (a combination of fluorine and oxygen); that is introduced into the second stage reaction chamber 17 by an oxidizer injector 24. The reaction occurring between the oxidizer and the products of decomposition of the hydrazine produces an elevated temperature in the second stage reaction chamber (e.g. approximately 5,0007,000F).

According to the invention a heat pipe HP is employed for removing some of the heat energy from the chamber walls so that conventional wall materials may be used. The heat pipe HP includes a tubular evaporator section 32 that may conform in shape to the outline of the rocket engine. It also includes a condenser section 34 which in the preferred embodiment is located between the oxidizer injector 24 and wall 18, within the flow path of the products of decomposition exiting from the first stage reaction chamber 16.

General heat pipe principles, as discussed in the aforementioned Scientific American article, are employed. The two sections 32 and 34 are joined to form a unitary evacuated chamber. Section 32 is defined by a casing 36 (preferably also made of stainless steel or the like). A wick 38 is housed within casing 36 and is shown to be in contact with the inner wall of the casing 36 which in turn is in contact with the portion of thrust chamber wall 12 which surrounds the second stage reaction chamber 17.

Preferably, the wick is formed of laminates of stainless steel screen wire (120 X 120 mesh, for example) that are spot welded together into three or four layers. However, as discussed in the aforementioned article, any material that provides sufficient numbers of small passages to provide efficient capillary action and which is compatible with the coolant is sufficient.

The heat pipe working fluid may be a liquid metal such as lithium, which has an evaporization temperature of 2,180F at 5 psia. The vaporization temperature may be varied by varying the partial vacuum pressure in the heat pipe HP. The liquid coolant carried in the wick travels toward the nozzle end of the rocket engine and is vaporized by the heat transferred from the high temperature combustion gases. The coolant vapor then flows in the direction of the arrows 40 toward the condenser section of the heat pipe HP.

The preferred form of the condenser section 34 of the heat pipe HP is best shown in FIGS. 2 and 3. It is a tubetype heat exchanger comprising a plurality of parallel, spaced apart tubes 42 each one of which is a continuation of the heat pipe. By way of typical example, for a 1,000 pound thrust engine having a chamber pressure of approximately psia and a combustion temperature of approximately 5,510F. a seven row, seven tube per row heat exchanger is theoretically calculated to be preferred. The hydrazone flow rate for these parameters is approximately 1.3 pounds per second with an oxidizer (nitrogen tetroxide) flow rate of approximately 1.8 pounds per second. The temperature rise of the products of decomposition through the heat exchanger is approximately 180F. and it experiences a pressure drop of approximately 14.5 psi. The coolant flow rate (lithium) is approximately 0.0175 pounds per second and the heat transfer rate to the heat pipe at the chamber walls is approximately 147 btu per second. The calculated heat exchanger conductance of the preferred embodiment is 0.307 btu per second degrees Rankin.

It is necessary, of course, to connect the wick 38 in the condenser section 34 with the wick in the vaporization section 32. As best shown in FIG. 3, this is accomplished by welding the tubes of the heat exchanger to the thrust chamber wall 12. The wick thus provides for a liquid path and the open space around the wick forms a vapor path. The vaporized coolant is condensed to the liquid state in heat exchanger 42 and by capillary action passes through the wick 38 into the vaporization section.

In the embodiment shown in FIG. 4, the products of hydrazine decomposition are again used to condense the coolant in the heat pipe HP. One of the difficulties of the embodiment of FIGS. 1-3 is that the high pressure in the thrust chamber may cause leakage into the heat pipe casing unless the welding or other connec tions around the heat exchanger tubes are adequate. In the form of engine shown in FIG. 4, however, this difficulty is eliminated by providing the heat exchanger externally of the thrust chamber. Suitable ducting 44 is employed to remove a portion of the products of decomposition and direct it into a heat exchanger 46 which is in heat transfer relationship with the condenser section 48 of the heat pipe I-IP. In this form the vapors which leave the inner wick zone 38 surrounding the nozzle 14' condense into the outer wick zone 38' positioned next to passageway 46. In this form of engine propulsive gases which are removed for use as a condensing medium are not available as a propellant to react with the liquid oxidizer. However, the removed gases may be delivered through nozzles N for use as monopropellant thrusters to supplement the thrust produced in the main engine. Of course, other applications may also be found for the propulsive gases after they leave the heat exchanger.

In the embodiment of FIG. 4 the bipropellant reaction region or zone 47 is surrouneded by annular zone 49 of decomposed gases from the catalyst bed 16. Although not shown, these same zones 47, 49 are formed during operation of the embodiment shown by FIGS. 1-3. The relatively cooler outer zone 49 lessens the cooling problem by partially protecting the chamber walls from the relatively high temperature gases in zone 47.

Other forms of removing the heat energy at the condenser section 48, not shown, include (I) use of one or both of the liquid propellants to absorb heat in the condenser section 48 prior to entering injector 20 and/or 24; and (2) providing sufficient surface area of the condenser section 48 to radiate the thermal energy to the ambient environment.

In the embodiment shown in FIG. the heat exchanger is supplied through line 50 with a liquid propellant P or P as the condensing medium. The

liquid propellant P or P may be one of the two propellants of a bipropellant rocket or it may be the liquid propellant used with the second thrust level stage of an engine such as in the embodiments of FIGS. 1-4. In this embodiment the absorbing wick region 52 is shown located upstream of the injector head 54, as a colinear extension of the wick region 56 which immediately surrounds the reaction chamber wall 58 and nozzle 60.

The propellants are injected in the combustion chamber with a fraction of one of the propellants being injected along the chamber wall 58, thereby providing a barrier to reduce heat transfer from the combustion gases to the heat pipe wick 56. The propellant is delivered as a liquid film or sheath 64 along the inner surface of chamber wall 56. It is soon vaporized and becomes a vaporous film 66 shown to continue flowing along walls 58 and 60.

What is claimed is:

1. A rocket engine having a thrust chamber comprising:

a primary reaction zone within said chamber having a propellant which forms primary propulsive gases at a relatively low temperature;

a secondary reaction zone in the path of said primary discharge gases;

an oxidizer inlet in said path for introducing an oxidizer into said primary discharge gases, said oxidizer and said primary discharge gases reacting to form secondary propulsive gases at a temperature substantially greater than said primary propulsive gases;

means for cooling the thrust chamber walls surrounding said secondary reaction zone, said means including a liquid coolant; and

heat exchanger means disposed to be contacted by at least a portion of said primary propulsive gases and communicating with said cooling means for receiving said coolant, whereby the relatively low temperature primary propulsive gases are used to cool the coolant.

2. The rocket engine defined by claim 1, wherein said cooling means includes an evacuated heat pipe having a wick and said coolant includes a fluid that is a liquid at the temperature of said primary propulsive gases and vaporizes at the temperature of said secondary propulsive gases.

3. The rocket engine defined by claim 2, wherein said coolant includes lithium and said propellant includes hydrazine.

4. The rocket engine defined by claim 3, wherein said oxidizer is nitrogen tetroxide, fluorine, oxygen, or combinations of fluorine and oxygen.

5. The rocket engine defined by claim 1, wherein said primary propulsive gases include the products ofa catalytically decomposable monopropellant.

6. The rocket engine defined by claim 2, wherein said heat exchanger means includes a plurality of tubes disposed within said thrust chamber each tube containing an extension of said wick and an open vapor channel.

7. The rocket engine defined by claim 2, wherein said heat exchanger is located externally of said thrust chamber and further including duct means for directing at least a portion of said primary propulsive gases out of said thrust chamber.

6 8. The rocket defined by claim 7, wherein said pripassing the coolant in indirect heat exchange relamary propulsive gases are passed through a tionship with the article to be cooled; monopropellant thruster after leaving said h at then passing the heated coolant in indirect heat exchanger. exchange relationship with at least a portion of the 9- A method of cooling the Chamber Walls Of a 5 decomposition products of said fuel component, bipropellant engine comprising? so that the latter can receive heat from and thus generating a relatively low temperature, low thrust h f me d fuel gas Within a reaction chamber; delivering the heated decomposition products of said introducing an oxidizer into the path of said fuel gas fuel component into a combustion chamber f for combustion into a higher temperature, higher 10 thrust propulsive gas and as a result heating the chamber walls;

cooling said chamber walls by circulating a coolant there around; and

cooling said coolant by passing the coolant through said propulsive fuel gas upstream of said oxidizer.

10. The method defined by claim 9, wherein said step of generating fuel gas within said chamber includes catalytically decomposing hydrazine.

11. The method of claim 9, wherein said coolant is passed indirectly through said propulsive fuel gas within said chamber.

12. In a system wherein a chemical fuel component thfereabfmt; and of fluid form is consumed and a heated article is cooled coolfng coolant by d'rectmg a econd porno of by arecirculated fluid coolant, the method comprising: gas out from the chamber and decomposing the fuel component to form decom intorndirect heat exchange with said coolant.

position products;

same while at the same time recirculating the now cooled coolant back into indirect heat relationship with said article.

13. A method of cooling the chamber walls of a bipropellant engine comprising:

generating a relatively low temperature, low thrust fuel gas within a reaction chamber;

introducing an oxidizer into the path ofa first portion of said fuel gas for combustion into a higher temperature, higher thrust propulsive gas and as a result heating the chamber walls;

cooling said chamber walls by circulating a coolant

Claims (12)

1. A rocket engine having a thrust chamber comprising: a primary reaction zone within said chamber having a propellant which forms primary propulsive gases at a reLatively low temperature; a secondary reaction zone in the path of said primary discharge gases; an oxidizer inlet in said path for introducing an oxidizer into said primary discharge gases, said oxidizer and said primary discharge gases reacting to form secondary propulsive gases at a temperature substantially greater than said primary propulsive gases; means for cooling the thrust chamber walls surrounding said secondary reaction zone, said means including a liquid coolant; and heat exchanger means disposed to be contacted by at least a portion of said primary propulsive gases and communicating with said cooling means for receiving said coolant, whereby the relatively low temperature primary propulsive gases are used to cool the coolant.
2. The rocket engine defined by claim 1, wherein said cooling means includes an evacuated heat pipe having a wick and said coolant includes a fluid that is a liquid at the temperature of said primary propulsive gases and vaporizes at the temperature of said secondary propulsive gases.
3. The rocket engine defined by claim 2, wherein said coolant includes lithium and said propellant includes hydrazine.
4. The rocket engine defined by claim 3, wherein said oxidizer is nitrogen tetroxide, fluorine, oxygen, or combinations of fluorine and oxygen.
5. The rocket engine defined by claim 1, wherein said primary propulsive gases include the products of a catalytically decomposable monopropellant.
6. The rocket engine defined by claim 2, wherein said heat exchanger means includes a plurality of tubes disposed within said thrust chamber each tube containing an extension of said wick and an open vapor channel.
7. The rocket engine defined by claim 2, wherein said heat exchanger is located externally of said thrust chamber and further including duct means for directing at least a portion of said primary propulsive gases out of said thrust chamber.
8. The rocket defined by claim 7, wherein said primary propulsive gases are passed through a monopropellant thruster after leaving said heat exchanger.
9. A method of cooling the chamber walls of a bipropellant engine comprising: generating a relatively low temperature, low thrust fuel gas within a reaction chamber; introducing an oxidizer into the path of said fuel gas for combustion into a higher temperature, higher thrust propulsive gas and as a result heating the chamber walls; cooling said chamber walls by circulating a coolant there around; and cooling said coolant by passing the coolant through said propulsive fuel gas upstream of said oxidizer.
10. The method defined by claim 9, wherein said step of generating fuel gas within said chamber includes catalytically decomposing hydrazine.
11. The method of claim 9, wherein said coolant is passed indirectly through said propulsive fuel gas within said chamber.
12. In a system wherein a chemical fuel component of fluid form is consumed and a heated article is cooled by a recirculated fluid coolant, the method comprising: decomposing the fuel component to form decomposition products; passing the coolant in indirect heat exchange relationship with the article to be cooled; then passing the heated coolant in indirect heat exchange relationship with at least a portion of the decomposition products of said fuel component, so that the latter can receive heat from and thus cool the former; and delivering the heated decomposition products of said fuel component into a combustion chamber for same while at the same time recirculating the now cooled coolant back into indirect heat relationship with said article.
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Cited By (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3871828A (en) * 1972-10-10 1975-03-18 Hughes Aircraft Co Hydrazine gas generator
US3893294A (en) * 1973-09-10 1975-07-08 United Aircraft Corp Catalytic monopropellant reactor with thermal feedback
US4199937A (en) * 1975-03-19 1980-04-29 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Heat exchanger and method of making
US4245469A (en) * 1979-04-23 1981-01-20 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Heat exchanger and method of making
US4324096A (en) * 1979-04-23 1982-04-13 Hughes Aircraft Company Hydrazine thruster
US4406863A (en) * 1982-02-09 1983-09-27 The United States Of America As Represented By The Secretary Of The Air Force Integrated solid propellant gas generator and fluid heat exchanger
US4470258A (en) * 1981-07-21 1984-09-11 Erno Raumfahrttechnik Gmbh Thruster for space vehicles
FR2556046A1 (en) * 1983-12-02 1985-06-07 United Technologies Corp heater protection for impellers
US4578946A (en) * 1984-09-17 1986-04-01 Sundstrand Corporation Hydrazine fuel catalytic gas generator and injector therefor
US5442910A (en) * 1994-03-21 1995-08-22 Thermacore, Inc. Reaction motor structure and method of construction
WO1996012688A1 (en) * 1994-10-20 1996-05-02 Kunkel, Klaus Process for operating a reaction-type missile propulsion system and missile propulsion system
US5787702A (en) * 1995-06-13 1998-08-04 Daimler-Benz Aerospace Ag Propulsion plant operating on the basis of catalytic and/or chemical decomposition of a propellant
WO2000079115A1 (en) * 1999-06-17 2000-12-28 Astrium Gmbh Thrust chamber assembly
US20060243361A1 (en) * 2005-04-29 2006-11-02 Foxconn Technology Co., Ltd. Ageing process for sealed product
US20080173020A1 (en) * 2006-12-04 2008-07-24 Firestar Engineering, Llc Spark-integrated propellant injector head with flashback barrier
US20090133788A1 (en) * 2007-11-09 2009-05-28 Firestar Engineering, Llc Nitrous oxide fuel blend monopropellants
US20090211228A1 (en) * 2007-03-12 2009-08-27 Honeywell International, Inc. High performance liquid fuel combustion gas generator
US20100275577A1 (en) * 2006-12-04 2010-11-04 Firestar Engineering, Llc Rocket engine injectorhead with flashback barrier
US20110005195A1 (en) * 2009-07-07 2011-01-13 Firestar Engineering, Llc Aluminum porous media
US20110180032A1 (en) * 2010-01-20 2011-07-28 Firestar Engineering, Llc Insulated combustion chamber
US20110219742A1 (en) * 2010-03-12 2011-09-15 Firestar Engineering, Llc Supersonic combustor rocket nozzle
US20130199155A1 (en) * 2012-01-02 2013-08-08 Jordin Kare Rocket Propulsion Systems, and Related Methods
US8572946B2 (en) 2006-12-04 2013-11-05 Firestar Engineering, Llc Microfluidic flame barrier
US20140182265A1 (en) * 2013-01-03 2014-07-03 Jordin Kare Rocket Propulsion Systems, and Related Methods
GB2518211A (en) * 2013-09-13 2015-03-18 Carolyn Billie Knight Evaporative wick/membrane rocket motor
RU2563114C1 (en) * 2014-05-19 2015-09-20 Оао "Кузнецов" Liquid propellant rocket engine chamber nozzle
WO2016167700A1 (en) * 2015-04-14 2016-10-20 Ecaps Aktiebolag Liquid propellant chemical rocket engine reactor thermal management system

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US3232048A (en) * 1959-12-12 1966-02-01 Bolkow Gmbh Rocket engine
US3493177A (en) * 1967-07-26 1970-02-03 Trw Inc Method of and means for cooling the throat wall of rocket engine nozzle

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US3024606A (en) * 1958-07-10 1962-03-13 Curtiss Wright Corp Liquid cooling system for jet engines
US3232048A (en) * 1959-12-12 1966-02-01 Bolkow Gmbh Rocket engine
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US3493177A (en) * 1967-07-26 1970-02-03 Trw Inc Method of and means for cooling the throat wall of rocket engine nozzle

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3871828A (en) * 1972-10-10 1975-03-18 Hughes Aircraft Co Hydrazine gas generator
US3893294A (en) * 1973-09-10 1975-07-08 United Aircraft Corp Catalytic monopropellant reactor with thermal feedback
US4199937A (en) * 1975-03-19 1980-04-29 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Heat exchanger and method of making
US4245469A (en) * 1979-04-23 1981-01-20 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Heat exchanger and method of making
US4324096A (en) * 1979-04-23 1982-04-13 Hughes Aircraft Company Hydrazine thruster
US4470258A (en) * 1981-07-21 1984-09-11 Erno Raumfahrttechnik Gmbh Thruster for space vehicles
US4406863A (en) * 1982-02-09 1983-09-27 The United States Of America As Represented By The Secretary Of The Air Force Integrated solid propellant gas generator and fluid heat exchanger
US4583361A (en) * 1983-12-02 1986-04-22 United Technologies Corporation Heater protection of thrusters
FR2556046A1 (en) * 1983-12-02 1985-06-07 United Technologies Corp heater protection for impellers
US4578946A (en) * 1984-09-17 1986-04-01 Sundstrand Corporation Hydrazine fuel catalytic gas generator and injector therefor
US5442910A (en) * 1994-03-21 1995-08-22 Thermacore, Inc. Reaction motor structure and method of construction
US5579576A (en) * 1994-03-21 1996-12-03 Thermacore, Inc. Reaction motor structure and method of construction
WO1996012688A1 (en) * 1994-10-20 1996-05-02 Kunkel, Klaus Process for operating a reaction-type missile propulsion system and missile propulsion system
US5787702A (en) * 1995-06-13 1998-08-04 Daimler-Benz Aerospace Ag Propulsion plant operating on the basis of catalytic and/or chemical decomposition of a propellant
WO2000079115A1 (en) * 1999-06-17 2000-12-28 Astrium Gmbh Thrust chamber assembly
US6698184B1 (en) 1999-06-17 2004-03-02 Astrium Gmbh Thrust chamber assembly
US20060243361A1 (en) * 2005-04-29 2006-11-02 Foxconn Technology Co., Ltd. Ageing process for sealed product
US8230673B2 (en) * 2006-12-04 2012-07-31 Firestar Engineering, Llc Rocket engine injectorhead with flashback barrier
US20080173020A1 (en) * 2006-12-04 2008-07-24 Firestar Engineering, Llc Spark-integrated propellant injector head with flashback barrier
US8572946B2 (en) 2006-12-04 2013-11-05 Firestar Engineering, Llc Microfluidic flame barrier
US20100275577A1 (en) * 2006-12-04 2010-11-04 Firestar Engineering, Llc Rocket engine injectorhead with flashback barrier
US8230672B2 (en) * 2006-12-04 2012-07-31 Firestar Engineering, Llc Spark-integrated propellant injector head with flashback barrier
US20090211228A1 (en) * 2007-03-12 2009-08-27 Honeywell International, Inc. High performance liquid fuel combustion gas generator
US20090133788A1 (en) * 2007-11-09 2009-05-28 Firestar Engineering, Llc Nitrous oxide fuel blend monopropellants
US8858224B2 (en) 2009-07-07 2014-10-14 Firestar Engineering, Llc Detonation wave arrestor
US20110146231A1 (en) * 2009-07-07 2011-06-23 Firestar Engineering, Llc Tiered Porosity Flashback Suppressing Elements for Monopropellant or Pre-Mixed Bipropellant Systems
US20110008739A1 (en) * 2009-07-07 2011-01-13 Firestar Engineering, Llc Detonation wave arrestor
US20110005195A1 (en) * 2009-07-07 2011-01-13 Firestar Engineering, Llc Aluminum porous media
US20110180032A1 (en) * 2010-01-20 2011-07-28 Firestar Engineering, Llc Insulated combustion chamber
US20110219742A1 (en) * 2010-03-12 2011-09-15 Firestar Engineering, Llc Supersonic combustor rocket nozzle
US20130199155A1 (en) * 2012-01-02 2013-08-08 Jordin Kare Rocket Propulsion Systems, and Related Methods
US20140182265A1 (en) * 2013-01-03 2014-07-03 Jordin Kare Rocket Propulsion Systems, and Related Methods
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
RU2563114C1 (en) * 2014-05-19 2015-09-20 Оао "Кузнецов" Liquid propellant rocket engine chamber nozzle
WO2016167700A1 (en) * 2015-04-14 2016-10-20 Ecaps Aktiebolag Liquid propellant chemical rocket engine reactor thermal management system

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