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Publication numberUS3695041 A
Publication typeGrant
Publication dateOct 3, 1972
Filing dateMay 8, 1970
Priority dateMay 8, 1970
Also published asDE2122742A1
Publication numberUS 3695041 A, US 3695041A, US-A-3695041, US3695041 A, US3695041A
InventorsEggers Robert F, Emmons Donald L
Original AssigneeRocket Research Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Two-stage hydrazine rocket motor
US 3695041 A
Abstract
Hydrazine is catalytically decomposed in a primary reaction chamber. During low thrust operation all of the decomposition products flow directly to the nozzle. During high thrust operation an oxidizer is reacted with a major portion of the decomposed hydrazine in the central portion of the secondary reaction chamber downstream of the primary reaction chamber. The remaining decomposed hydrazine is directed as a buffer layer along the inner surface of the secondary reaction chamber's outer wall, to serve as a coolant.
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United States Patent [151 3,695,041

Eggers et a1. 1 Oct. 3, 1972 [54] TWO-STAGE HYDRAZINE ROCKET 2,706,887 4/1955 Gron ..60/258 MOTOR 3,107,485 10/1963 Toulmin ..60/260 3 377 140 4/1968 Hall ..60/258 [72] Inventors. Robert F. Eggers Snohomtsh,

L. Emmons q MOSlel' f W h. o as Primary Examiner-Mark M. Newman Asslgneei f' Research Corporation, Assistant Examiner-Warren Olsen Rlchmond wash- Attorney-Graybeal, Cole & Barnard [22] Filed: May 8, 1970 [57] ABSTRACT [21] Appl. No.: 35,858

Hydrazme 1s catalytlcally decomposed m a primary reaction chamber. During low thrust operation all of 318.81 60/26:}, 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. 3,029,602 4/1962 Allen ..60/260 3,149,460 9/1964 La Rocca ..'....60/260 8 Claims, 3 Drawing Figures TWO-STAGE HYDRAZINE ROCKET MOTOR BACKGROUND OF THE INVENTION 1 Field of the Invention 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.

2. Description of the Prior Art Hydrazine in bipropellant rocket engines has been considered an advantageous fuel because of its high performance and its favorable storage and handling characteristics. The primary difficulties heretofore encountered in the use of liquid hydrazine in bipropellant engines are that the explosive thermal decomposition property of hydrazine has seriously limited its use as a reliable regenerative coolant and that the liquid injection of the propellants, such as hydrazine and nitrogen tetroxide frequently result in rough starting transients and combustion instability. One technique of improving the ignition characteristics of these propellants has been to substitute for the hydrazine another propellant, such as Aerozine 50 (N H /UDMH) or MMH. Unfortunately, however, these hydrazine-substitutes produce lower specific impulse than hydrazine. In addition, 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.

- Another technique for using hydrazine as the main fuel has been to decompose the hydrazine to produce first stage thrust and to introduce an oxidizer into the products of decomposition to produce a higher secondary stage thrust. In this case the products of decomposition, N N NH readily react with the oxidizer, N or F for example, to produce the desired high performance and thus eliminate the use of a hydrazinesubstitute and its storage facility. Such a technique is shown, for example, in the United States patent to LaRocca, US. Pat. No. 3,149,460, granted Sept. 22, I964.

In the reaction between the hydrazine decomposition products and an oxidizer an enormous quantity of heat is generated. 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.

One technique for cooling reaction chamber walls is shown in the United States patent to Grow, US. Pat. No. 2,706,887, granted Apr. 26, 1955. In this patent an annular liner is placed in the reaction chamber and one of the propellants is introduced inside and outside of the liner. The propellant introduced inside the liner is reacted with a second propellant producing high temperature reaction products. The propellant introduced outside the annular shield is decomposed in such outer chamber to produce a lower temperature motive fluid.

SUMMARY OF THE INVENTION 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.

According to the invention, 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.

BRIEF DESCRIPTION OF THE DRAWING 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; and

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.

DESCRIPTION OF THE PREFERRED EMBODIMENTS 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. In the illustrated embodiment 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.

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. In FIGS. 1 and 2 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. In a linerless embodiment the outlet ports 39 would be directed generally axially.

In operation, 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. During single stage operation all of the decomposed hydrazine flows directly to the nozzle 20. When higher thrust capability is required, the gases exiting the openings 33 are mixed with the oxidizer (e.g., nitrogen tetroxide (N 0,) which is introduced through the injector 37. 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.

During operation, the pressure drop of the gases within chamber 14 exerts a force which tends to buckle wall 32 outwardly. Preferably the conduits 24 are anchored at their ends so that they can provide support for wall 32.

in the embodiment shown in FIG. 3 the basic structural configuration is the same as that shown in FIG. 1. However, 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.

Many space vehicles have propulsion requirements that require an engine to function not only in steady state, but also in pulse mode operation. Delivered specific impulse of conventional bipropellant engines is generally low during pulse mode operation especially for ,small impulse bits. Using the dual mode feature of the two-stage hydrazine rocket motor concept, the thruster can be operated as a monopropellant engine during pulse mode, and as a bipropellant during steady state; thus delivering relatively high specific impulse during both modes of operation. The Biplex system is significantly lighter than the conventional bipropellant system, especially for missions which require small impusle bits during pulse mode operation.

Principal advantages of the engine of this invention include:

1. 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. 2. 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. System flexibility can be achieved through use of common hydrazine propellant tanks for both the main engine and attitude control system. This is attractive from a system viewpoint since significant deviations from a nominal mission can be accommodated by changing the propellant allocations for the attitude control system and the main engine.

While two forms of the invention have been illustrated various other modifications will be apparent to those skilled in the art. The details disclosed therefor are not to be construed as limitations on the invention.

What is claimed is: i

l. A gas generator comprising:

a primary reaction chamber;

a secondary reaction chamber downstream of said primary reaction chamber defined by wall mean comprising at least one tubular wall; I

means for introducing a primary reactant into said primary reaction chamber;

means in said primary reaction chamber for reacting said primary reactant to produce relatively low temperature but combustible discharge products;

means for introducing a first portion of said discharge products into an inner region of said secondary reaction chamber to serve as a reactant, and a' second portion as a continuous annular stream along said tubular wall, to function as a coolant; and

means for introducing a second reactant into said inner region of said secondary reaction chamber, in a zone laterally inwardly of the coolant, for reacting with thedischarge products in such zone to cause a higher temperature reaction.

2. The gas generator defined by claim 1, wherein said reacting means in said primary reaction chamber includes a catalyst and said primary reactant comprises hydrazine. I

3. The gas generator defined by claim 1, wherein said second reactant is an oxidizer.

4. The gas generator defined by claim 1, wherein in addition to said 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; and

an annular liner in said secondary reaction chamber separating the higher temperature reaction zone from the chamber wall coolant; and

wherein 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.

6. The gas generator defined by claim 5, wherein the means for introducing a second reactant comprises an annular manifold inside said secondary reaction chamber and said pervious wall includes openings directed to introduce discharge products through said manifold and openings directed to introduce discharge products into the space around said manifold, and said-

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2706887 *Jan 23, 1946Apr 26, 1955Grow Harlow BLiquid propellant rocket motor
US3029602 *Jun 16, 1958Apr 17, 1962Bristol Siddeley Engines LtdCombustion chambers
US3107485 *May 27, 1959Oct 22, 1963Ohio Commw Eng CoPropulsion means and method for space vehicles employing a volatile alkene and metalcarbonyl
US3149460 *Sep 28, 1960Sep 22, 1964Gen ElectricReaction propulsion system
US3377140 *Oct 15, 1965Apr 9, 1968Specialties Dev CorpApparatus for catalytically decomposing hydrazine
US3446023 *Aug 5, 1966May 27, 1969United Aircraft CorpCatalytic attitude-control rocket motor
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3871828 *Oct 10, 1972Mar 18, 1975Hughes Aircraft CoHydrazine gas generator
US3903693 *Mar 26, 1973Sep 9, 1975Fox AnthonyRocket motor housing
US4069664 *Mar 22, 1976Jan 24, 1978Hughes Aircraft CompanyMonopropellant thruster
US4324096 *Apr 23, 1979Apr 13, 1982Hughes Aircraft CompanyHydrazine thruster
US4856271 *Oct 1, 1987Aug 15, 1989Olin CorporationGas generator and generating method employing dual catalytic and thermal liquid propellant decomposition paths
US5711695 *May 1, 1995Jan 27, 1998Pitsco, Inc.Gas-propelled toy with exhaust nozzle for gas cartridge
US6860099 *Jan 9, 2003Mar 1, 2005The United States Of America As Represented By The Administrator Of The National Aeronautics And Space AdministrationLiquid propellant tracing impingement injector
US6918243 *May 19, 2003Jul 19, 2005The Boeing CompanyBi-propellant injector with flame-holding zone igniter
US7827781 *Sep 26, 2006Nov 9, 2010Bendel Timothy BLiquid propellant rocket engine with pintle injector and acoustic dampening
US8122703Apr 28, 2006Feb 28, 2012United Technologies CorporationCoaxial ignition assembly
US20040231318 *May 19, 2003Nov 25, 2004Fisher Steven C.Bi-propellant injector with flame-holding zone igniter
US20130199155 *Jan 3, 2013Aug 8, 2013Jordin KareRocket Propulsion Systems, and Related Methods
US20140182265 *Jan 3, 2013Jul 3, 2014Jordin KareRocket Propulsion Systems, and Related Methods
Classifications
U.S. Classification60/264, 60/39.462, 60/258, 60/200.1
International ClassificationF02K9/00, C06B47/00, F02K9/08, C06B47/08
Cooperative ClassificationC06B47/08, F02K9/08
European ClassificationF02K9/08, C06B47/08