US3695041A - Two-stage hydrazine rocket motor - Google Patents

Two-stage hydrazine rocket motor Download PDF

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Publication number
US3695041A
US3695041A US35858A US3695041DA US3695041A US 3695041 A US3695041 A US 3695041A US 35858 A US35858 A US 35858A US 3695041D A US3695041D A US 3695041DA US 3695041 A US3695041 A US 3695041A
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Prior art keywords
reaction chamber
wall
primary
reactant
liner
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Expired - Lifetime
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US35858A
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English (en)
Inventor
Robert F Eggers
Donald L Emmons
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Rocket Research Co
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Rocket Research Co
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    • 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
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • C—CHEMISTRY; METALLURGY
    • C06—EXPLOSIVES; MATCHES
    • C06B—EXPLOSIVES OR THERMIC COMPOSITIONS; MANUFACTURE THEREOF; USE OF SINGLE SUBSTANCES AS EXPLOSIVES
    • C06B47/00—Compositions in which the components are separately stored until the moment of burning or explosion, e.g. "Sprengel"-type explosives; Suspensions of solid component in a normally non-explosive liquid phase, including a thickened aqueous phase
    • C06B47/02—Compositions in which the components are separately stored until the moment of burning or explosion, e.g. "Sprengel"-type explosives; Suspensions of solid component in a normally non-explosive liquid phase, including a thickened aqueous phase the components comprising a binary propellant
    • C06B47/08—Compositions in which the components are separately stored until the moment of burning or explosion, e.g. "Sprengel"-type explosives; Suspensions of solid component in a normally non-explosive liquid phase, including a thickened aqueous phase the components comprising a binary propellant a component containing hydrazine or a hydrazine derivative

Definitions

  • a the decomposition products flow directly to the n zlean ed Field Of Search a major portion of the decomposed hydrazine in 60/260 266 the central portion of the secondary reaction chamber downstream of the primary reaction chamber.
  • the [56] References and remaining decomposed hydrazine is directed as a UNITED STATES PATENTS buffer layer along the inner surface of the secondary reaction chambers outer wall, to serve as a coolant.
  • This invention relates to gas generators such as in rocket propulsion systems, and more particularly to a two-stage propulsion system which uses the reacted products of a monopropellant first stage as both a propellant and a coolant for a bipropellant second stage.
  • hydrazine-substitutes produce lower specific impulse than hydrazine.
  • hydrazine is already being more widely used as a monopropellant for attitude control and gas pressurization systems and therefore it is highly desirable to avoid using a hydrazine-substitute so as to eliminate the need for a third storage facility aboard the space vehicle.
  • the thrust chamber walls and nozzle in order to be protected must either be manufactured of heat resisting materials, lined with ceramic materials or the like, cooled by passing a coolant through or over them, or combinations or these.
  • This invention pertains to a two-stage rocket motor which employs a primary reaction chamber in which hydrazine is catalytically decomposed, and a secondary reaction chamber in which the hydrazine products of decomposition may be reacted with an oxidizer.
  • the products of hydrazine decomposition may be used alone to provide low energy propulsion and reacted with the oxidizer to provide high energy propulsion. Also, during the secondary reaction some of the products of hydrazine decomposition are advantageously used as a film coolant for the side wall of the secondary reaction chamber.
  • FIG. 1 is a view partially in side elevation and partially in longitudinal section of a two-stage rocket motor employing the principles of the invention
  • FIG. 2 is a transverse sectional view taken substantially along line 2-2 of FIG. 1, with a portion of the oxidizer injector cut away;
  • FIG. 3 is a fragmentary view, also partially in side elevation and partially in longitudinal section, of a modified form of two-stage rocket motor.
  • the illustrated embodiment comprises an outer tubular wall 12 of stainless steel or the like defining an internal space which is divided axially into a first stage or primary reaction chamber 14 and a second stage or secondary reaction chamber 16.
  • Wall 12 is provided with a peripheral flange 18 to which is secured a rocket nozzle 20, also of stainless steel or the like, which may be of conventional design.
  • a conduit 22 leads from a source of liquid hydrazine to a multiported injector incorporated into the upstream wall 26 of the primary reaction chamber.
  • a plurality of conduits 24 extend from a source of an oxidizer, such as nitrogen tetroxide, to and then axially through the primary reaction chamber 14.
  • the first stage reaction chamber 14 is divided into two parts by a perforated screen or plate 28.
  • the upstream part contains fine catalyst particles and the downstream part contains coarser catalyst particles.
  • Reaction chamber 14 has a perforated downstream wall 32.
  • a screen 30 is located inside of wall 32 for retaining the catalyst particles within chamber 14.
  • Wall 32 is provided with a plurality of central openings 31, a plurality of intermediate openings 33, and a plurality of outer openings 34.
  • the outer openings 34 are directed to discharge adjacent an inner surface portion of wall 12.
  • the second stage reaction chamber 16 may include an annular liner 36 (e.g., of a refractory material) which is spaced radially inwardly of outer wall 12, so as to leave an annular space between the liner 36 and the outer wall 12. In other embodiments (not shown) this liner may be omitted.
  • annular liner 36 e.g., of a refractory material
  • An annular oxidizer injector 37 is centrally positioned within the second stage reaction chamber, and is connected to the conduits 24 leading from the supply of liquid oxidizer.
  • the outlet ports 39 of injector 37 are shown to discharge radially outwardly into the annular zone surrounding injector 37, for mixing contact in such zone with hydrazine decomposition products.
  • the outlet ports 39 would be directed generally axially.
  • liquid hydrazine is delivered into the first stage reaction chamber 14 and decomposes therein, forming hot gaseous nitrogen, hydrogen and ammonia, the temperature of which is about of l,600 F.
  • the main portion of these gases passes through the openings 31, 33 in wall 32.
  • all of the decomposed hydrazine flows directly to the nozzle 20.
  • the gases exiting the openings 33 are mixed with the oxidizer (e.g., nitrogen tetroxide (N 0,) which is introduced through the injector 37.
  • N 0, nitrogen tetroxide
  • the second stage reaction between the hydrazine gases and the oxidizer results in an increase in the chamber gas temperature to approximately 5,000 F depending upon the oxidizer used.
  • the decomposed but relative cool (viz. about 1,600 F) hydrazine which flows through openings 34 in wall 32 into the space between the liner 36 and the wall 12 insulates the outer wall 12 and the nozzle 20 from the high temperature in the central zone.
  • the pressure drop of the gases within chamber 14 exerts a force which tends to buckle wall 32 outwardly.
  • the conduits 24 are anchored at their ends so that they can provide support for wall 32.
  • the basic structural configuration is the same as that shown in FIG. 1.
  • the oxidizer is directed through a conduit 24a which passes laterally through the outer chamber wall 12 and the liner 36, into an injector 37 spaced downstream of wall 32, and from injector 37 spaced downstream of wall 32, and from injector 37 into the secondary reaction chamber.
  • the start transient and combustion characteristics are inherently smoother due to the fact that only atomization and vaporization of the oxidizer are required for combustion to occur. This feature eliminates the problem of destructive ignition transients due to the accumulation of unburned or partially reacted fuel constituents.
  • the relatively cool decomposition products (approximately l,600 F) can be effectively used. for cooling the thrust chamber and thrust and total impulse levels. With valve modulation a greater throttling range with greater reliability can be obtained. 4. Due to the unique chamber and nozzle cooling capability, and the relatively simple injector elements, it is anticipated that low cost rocket engines in a wide range of thrusts can be tkveloped. 5.
  • a gas generator comprising:
  • a secondary reaction chamber downstream of said primary reaction chamber defined by wall mean comprising at least one tubular wall
  • said secondary reaction chamber wall means includes a tubular liner separating the higher temperature reaction zone from the chamber wall coolant, said liner is spaced radially inwardly from said tubular wall, and said annular coolant stream flows between such wall and said liner in conwith the discharge products in such zone to cause a higher temperature reaction;
  • the means for introducing said discharge products into said secondary reaction chamber comprises a pervious wall having first openings leading to the zone inside said liner and second openings leading to the zone surrounding said liner.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Organic Chemistry (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Organic Low-Molecular-Weight Compounds And Preparation Thereof (AREA)
US35858A 1970-05-08 1970-05-08 Two-stage hydrazine rocket motor Expired - Lifetime US3695041A (en)

Applications Claiming Priority (1)

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US3585870A 1970-05-08 1970-05-08

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DE (1) DE2122742A1 (enExample)
FR (1) FR2093469A5 (enExample)

Cited By (19)

* 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
US3903693A (en) * 1973-03-26 1975-09-09 Anthony Fox Rocket motor housing
US4069664A (en) * 1974-01-24 1978-01-24 Hughes Aircraft Company Monopropellant thruster
FR2455184A1 (fr) * 1979-04-23 1980-11-21 Hughes Aircraft Co Generateur de poussee a hydrazine et son procede de mise en oeuvre
US4856271A (en) * 1987-10-01 1989-08-15 Olin Corporation Gas generator and generating method employing dual catalytic and thermal liquid propellant decomposition paths
US5711695A (en) * 1995-05-01 1998-01-27 Pitsco, Inc. Gas-propelled toy with exhaust nozzle for gas cartridge
US20040231318A1 (en) * 2003-05-19 2004-11-25 Fisher Steven C. Bi-propellant injector with flame-holding zone igniter
US6860099B1 (en) * 2003-01-09 2005-03-01 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Liquid propellant tracing impingement injector
US20070062176A1 (en) * 2005-12-05 2007-03-22 Bendel Timothy B Liquid propellant rocket engine with pintle injector and acoustic dampening
US8122703B2 (en) 2006-04-28 2012-02-28 United Technologies Corporation Coaxial ignition assembly
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
US20220120240A1 (en) * 2020-10-16 2022-04-21 Sierra Nevada Corporation Vortex thruster system including catalyst bed with screen assembly
US11572851B2 (en) * 2019-06-21 2023-02-07 Sierra Space Corporation Reaction control vortex thruster system
US11661907B2 (en) 2018-10-11 2023-05-30 Sierra Space Corporation Vortex hybrid rocket motor
US20230193857A1 (en) * 2021-12-21 2023-06-22 Firehawk Aerospace, Inc. Catalytic decomposition reactors
US11879414B2 (en) 2022-04-12 2024-01-23 Sierra Space Corporation Hybrid rocket oxidizer flow control system including regression rate sensors
US11952967B2 (en) 2021-08-19 2024-04-09 Sierra Space Corporation Liquid propellant injector for vortex hybrid rocket motor
US11952965B2 (en) 2019-01-30 2024-04-09 Laboratoire Reaction Dynamics Inc. Rocket engine's thrust chamber assembly

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2706887A (en) * 1946-01-23 1955-04-26 Harlow B Grow Liquid propellant rocket motor
US3029602A (en) * 1957-06-21 1962-04-17 Bristol Siddeley Engines Ltd Combustion chambers
US3107485A (en) * 1959-05-27 1963-10-22 Ohio Commw Eng Co Propulsion means and method for space vehicles employing a volatile alkene and metalcarbonyl
US3149460A (en) * 1960-09-28 1964-09-22 Gen Electric Reaction propulsion system
US3377140A (en) * 1965-10-15 1968-04-09 Specialties Dev Corp Apparatus for catalytically decomposing hydrazine
US3446023A (en) * 1966-08-05 1969-05-27 United Aircraft Corp Catalytic attitude-control rocket motor

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2706887A (en) * 1946-01-23 1955-04-26 Harlow B Grow Liquid propellant rocket motor
US3029602A (en) * 1957-06-21 1962-04-17 Bristol Siddeley Engines Ltd Combustion chambers
US3107485A (en) * 1959-05-27 1963-10-22 Ohio Commw Eng Co Propulsion means and method for space vehicles employing a volatile alkene and metalcarbonyl
US3149460A (en) * 1960-09-28 1964-09-22 Gen Electric Reaction propulsion system
US3377140A (en) * 1965-10-15 1968-04-09 Specialties Dev Corp Apparatus for catalytically decomposing hydrazine
US3446023A (en) * 1966-08-05 1969-05-27 United Aircraft Corp Catalytic attitude-control rocket motor

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
US3903693A (en) * 1973-03-26 1975-09-09 Anthony Fox Rocket motor housing
US4069664A (en) * 1974-01-24 1978-01-24 Hughes Aircraft Company Monopropellant thruster
FR2455184A1 (fr) * 1979-04-23 1980-11-21 Hughes Aircraft Co Generateur de poussee a hydrazine et son procede de mise en oeuvre
US4324096A (en) * 1979-04-23 1982-04-13 Hughes Aircraft Company Hydrazine thruster
US4856271A (en) * 1987-10-01 1989-08-15 Olin Corporation Gas generator and generating method employing dual catalytic and thermal liquid propellant decomposition paths
US5711695A (en) * 1995-05-01 1998-01-27 Pitsco, Inc. Gas-propelled toy with exhaust nozzle for gas cartridge
US6860099B1 (en) * 2003-01-09 2005-03-01 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Liquid propellant tracing impingement injector
US20040231318A1 (en) * 2003-05-19 2004-11-25 Fisher Steven C. Bi-propellant injector with flame-holding zone igniter
US6918243B2 (en) * 2003-05-19 2005-07-19 The Boeing Company Bi-propellant injector with flame-holding zone igniter
US20070062176A1 (en) * 2005-12-05 2007-03-22 Bendel Timothy B Liquid propellant rocket engine with pintle injector and acoustic dampening
US7827781B2 (en) * 2005-12-05 2010-11-09 Bendel Timothy B Liquid propellant rocket engine with pintle injector and acoustic dampening
US8122703B2 (en) 2006-04-28 2012-02-28 United Technologies Corporation Coaxial ignition assembly
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
US11661907B2 (en) 2018-10-11 2023-05-30 Sierra Space Corporation Vortex hybrid rocket motor
US12071915B2 (en) 2018-10-11 2024-08-27 Sierra Space Corporation Vortex hybrid rocket motor
US11952965B2 (en) 2019-01-30 2024-04-09 Laboratoire Reaction Dynamics Inc. Rocket engine's thrust chamber assembly
US12129814B2 (en) 2019-01-30 2024-10-29 Laboratoire Reaction Dynamics Inc. Thrust vector control for hybrid propellants rocket engine with embedded fluid injection ports
US12060853B2 (en) 2019-01-30 2024-08-13 Laboratoire Reaction Dynamics Inc. Rocket engine with integrated oxidizer catalyst in manifold and injector assembly
US11572851B2 (en) * 2019-06-21 2023-02-07 Sierra Space Corporation Reaction control vortex thruster system
US11927152B2 (en) * 2019-06-21 2024-03-12 Sierra Space Corporation Reaction control vortex thruster system
US20220120240A1 (en) * 2020-10-16 2022-04-21 Sierra Nevada Corporation Vortex thruster system including catalyst bed with screen assembly
US11952967B2 (en) 2021-08-19 2024-04-09 Sierra Space Corporation Liquid propellant injector for vortex hybrid rocket motor
US20230193857A1 (en) * 2021-12-21 2023-06-22 Firehawk Aerospace, Inc. Catalytic decomposition reactors
US12410765B2 (en) * 2021-12-21 2025-09-09 Firehawk Aerospace, Inc. Catalytic decomposition reactors
US11879414B2 (en) 2022-04-12 2024-01-23 Sierra Space Corporation Hybrid rocket oxidizer flow control system including regression rate sensors

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FR2093469A5 (enExample) 1972-01-28
DE2122742A1 (de) 1971-12-02

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