US3295321A - Method and apparatus for injecting a secondary propellant in multi-heat release combustors - Google Patents

Description

L. J. KRZYCKI METHOD AND APPARATUS FOR INJECTING A SECONDARY PROPELLANT IN MULTI-HEAT RELEASE GOMBUSTORS Jan. 3, 1967 Filed June 25, 1964 /l//ll///l//ll'l/ /lL//l//l/l FIG A (PRIOR ART) FIG B (PRIOR ART) C IO 1// ///l////////// ///l//l//l//l/l A FIG. L

lOa

LE ROY J.

INVENTOR.

BY V. C. MU L L E R ATTORN E Y.

FIG. 2.

United States Patent O 3,295,321 METHOD AND APPARATUS FOR INJECTING A SECONDARY PROPELLANT IN MULTl-HEAT RELEASE COMBUSTORS Leroy J. Krzycki, China Lake, Calif., assignor to the United States of America as represented by the Secretary of the Navy Filed June 25, 1964, Ser. No. 379,068 3 Claims. (Cl. 150-207) The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

This invention relates to rocket motors and more particularly to improvements in method and apparatus for injecting a liquid propellant.

In certain types of rockets it has been the practice to provide a multi-heat release combustor in which a primary non-stoichiometric reactor or heat release occurs in one zone of a combustion chamber and secondary reaction or heat release occurs downstream from the primary reaction in another zone as a result of injecting a secondary propellant, the latter being the propellant which was deficient in the non-stoichiometric primary reaction. In one of such types, known as peripheral injection, as illustrated in FIG. A, a liquid propellant P is injected through apertures A in the periphery of the combustion chamber at some locus downstream from the primary zone C but Upstream from the secondary heat release zone D. This system is undu ly complicated and bulky since a rnanifold is required for the secondary injection orifices, which becomes further complicated if the combustion chamber is provided with a surrounding coolant jacket. In another type, known as central pintle injection, as illustrated in FIG. B, a pintle or tube T is disposed axially of the combustion chamber, supported at its forward end by the forward closure of the combustion chamber and through the forward end of which the secondary propellant P is delivered. While this type obviates the disadvantages of peripheral injection, it has been found diflicult to maintain the structural integrity of the pintle in the presence of the hot erosive gases.

One of the objects of this invention is to provide novel method and apparatus for injecting a secondary propellant which obviates the disadvantages of the prior art devices referred to.

Another object is to eifect secondary injection with simplified apparatus.

A further object is to efiect secondary injection with the same apparatus employed for primary injection.

Still further objects, advantages and salient features Will become more apparent from a consideration of the description to follow, the appended claims, and the accompanying drawings in which:

FIGS. A and B are longitudinal central sections through rocket motors of the prior art.

FIG. 1 is a longitudial central section through a portion of a rocket motor, illustrating the invention in its broader aspects, and

FIG. 2 is a like section illustrating a specific embodiment of the invention.

The invention evolved from the testing of a pintle type secondary injector of the type illustrated in FIG. B. During this test, in an oxygen rich environment, the pintle wall eroded completely away, leaving only a relatively short tube projecting into the combustion chamber. It was noted, however, that the combustion performance remained satisfactory. From this discovery, an experiment was performed in which the pintle was shortened so that its discharge orifice Was fiush with the forward end of the combustion chamber. This was also found to perform satisfactorily. It became apparent, then, that the fuel stream of the secondary propellant could be injected through the primary heat release zone C with substantially no reaction with the oxygen rich gases in such zone. The reason for such -phenomena appears to be that the main mass of secondary injectant will not react with the hot flow of gases surrounding it since combustion normally occurs with respect to small droplets or vapor and not with respect to a large homogeneous or coherent mass of liquid. As the injectant traverses the primary heat release region a portion of the stream is sheared from the main stream due to two main influences. One of these is the relative velocity between the injected stream and the combustion gases surrounding it. The other is the density of the combustion gases. The coherent stream is surrounded by a mist of small droplets due to the shearing action and the jet breakup which is controlled by the turbulence and velocity profile of the jet and the density of the gases through which the jet is passing.

Through experiment with different size injectant orifices and injectant Velocities it was found that some of the injectant could be released from its coherent stream in the primary heat release Zone and the remainder delivered downstream to a secondary heat release zone. As will subsequently appear, the secondary heat release zone may be ahead of an expanding nozzle within the combustion chamber where the gas velocity is sub-sonic or rearward of the nozzle throat where the velocity is supersonic.

For Simplicity of disclosure, it will first be assumed that the coherent streams is to supply the requisite proportion of propellant for release in a primary heat release zone and the remainder for release in a secondary heat release zone, the secondary release zone being within the combustion chamber. It will also be assumed that the other propellant, either a gas or liquid, is introduced into the primary zone in any conventional manner (not shown); further, that the reaction within the primary release zone will be non-stoichiometric, producing combustion gases which are either oxygen-rich or fuel-rich and which can further react with the coherent stream injectant to provide an ultimate stoichiometric reaction.

Referring now to FIG. 1, combustion chamber 10 may be of any conventional form, such as an elongated cylindrical tube, having an exhaust nozzle 12 as its rear end which will maintain desired combustion chamber pressure and expand the exhaust gases to high velocity, producng thrust. At the forward end of the chamber an orifice 14 is provided for directing a coherent stream of liquid injectant 16 into the chamber along its longitudinal aXis. The characteristics of this stream will be assumed to be such that some of the injectant is released from the coherent stream for reaction with the other propellant in a primary heat release zone C and the remainder is released for subsequent reaction in a secondary release zone D.

The explanation of the control of disintegration of the coherent jet, to render it available at desired spaced heat release Zones, is not without its diflicultes; however, considerable investigations have been made of the phenomena. It is known, first, that the potential energy of the injected liquid is converted into kinetic energy in the discharge orifice. When the liquid leaves the orifice each particle may have both aXial and radial velocities. If the flow in the orifice is laminar and has a uniform velocity profile, there will be only an axial velocity. If the flow is turbulent, however, or if there is non-uniform velocity profile, a radial velocity Component will exist. During the flow through the orifice the liquid is restrained in the form of a continuous jet by the orifice wallwhich can introduce turbulence in the stream. As the jet leaves the orifice it begins to widen and disintegrate as the radial velocity and surrounding gas becomes effective. The length of the jet and its disintegration characteristics are thus dependent upon the design of the orifice. One of the major criteria of simple orifice injectors is the ration of orifice length to orifice diameter, the L/D ratio. At low values of L/D the eective jet, contracted at the entry to the orifice, has no time to re-expand and fill the passage. The velocity coemcient is compartively low. With increasing L/D the jet re-expands in the passage and the coeflicient increases, reaching a maximum for the simple orifice injector at a value of L/D between 4 and 6. Injection pressure has little eflect on the value of the coeflcent especially for low values of L/D (less than 3). For higher values of L/D, the ooefficient decreases with an increase in injection pressure because of increased jet contraction and increased velocity losses. Large L/D, however, allows sufficient passage length for the development of turbulent flow.

In the invention, as so far described, the secondary heat release zone is within the combustion chamber ahead of the nozzle. The process involved may be desired under some conditions, however, it suffers disadvantages of raising the temperature in the secondary release zone, and in some cases, to a temperature above that which the oombustion chamber will withstand with the particular propellants employed. To obviate this disadvantage, under such conditions, a specific application of the invention will now be described in which the secondary heat release zone is within the nozzle and rearwardly 'of the throat. By suitable choice of the fiow characteristics of the coherent injectant stream its final distintegration may be moved rearwardly so that the secondary heat release zone within the combustion chamber does not exist. As the combustion gases produced in the primary heat release zone move rearwardly they become enriched with particles from the coherent stream, but not sufiiciently so to spontaneously combust. As they move through the nozzle the mixture is now combustible but is precluded from reaction by reason of their supersonic velocity. It has been found, however, that such m'xture can be detonated within the divergent portion of a nozzle if triggered by a standing detonation wave. Such phenomena is fully described in the patent application of Clinton L. Spindler for Standing Detonation Wave Rocket Engine, Serial No. 31l,612, filed September 20, 1963.

Referring now to FIG. 2, a combustion chamber z is disclosed which is provided with a convergent-divergent de Laval nozzle 18 having two discrete divergent portions 20, 22 with a transition portion 24 between them. This discontinuity in the expansion cone has been found to produce a standing shock wave 26 which triggers the detonation in a secondary heat release zone 28 within the nozzle which is efiective to produce final stoichiometric reaction and produce additional thrust. This construction thus performs in a manner similar to that of the Spindler application with the exception that the pr-opellant required for the detonation in the secondary heat release Zone is supplied by the coherent stream, rather than by injecting into the nozzle throat. Also, the same stream Supplies the propellant for the primary heat release zone, rather than a separate injector. Thus, with a single injector, rather than with two, as in the Spindler application, two heat release zones are provided, one within the combustion chamber, and the other as a standing detonation wave within the expansion nozzle.

Obviously many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within the soope of the appended claims the invention may be practiced otherwise than as specifically described.

What is claimed is:

ll. A method of combusting bi-propellants, one of which is a liquid, in a rocket motor having a primary non-stoichiometric reaction zone and a spaced secondary stoichiometrc reaction zone downstream thereof within the divergent portion of a convergent-dvergent nozzle comprising the steps of:

(a) injecting said liquid from the forward end of said motor in a direction aXially thereof and parallel to the flow of combustion gases in the motor in the presence of the other propellant and in a coherent stream and at velocity such that a portion of the stream is removed therefrom for reaction in the primary zone,

(b) mixing the products of combustion from the primary zone with the remainder of said stream between the primary zone and the secondary zone, and

(c) combusting the last named mixture in the secondary zone.

2. A method of combusting bi-propellants in a rocket motor having a combustion chamber communicating with a supersonic convergent-divergent expansi-on nozzle, one of said propellants being a liquid, the combustion being non-stoichiometric in a primary Zone within the chamber and substantally stoichiometric in a secondary zone within the divergent portion of the nozzle rearwardly of a standing detonation wave in the divergent portion of the nozzle, comprising the steps of:

(a) injecting said liquid from the forward end of said chamber in a direction axially of said nozzle and parallel to the flow of combustion gases in the chamber in the presence of the other propellant and in a coherent stream and at a velocity such that a portion of the stream is removed therefrom for reaction in the primary zone,

(b) mixing the products of combustion from the primary zone with the remainder of said stream between the primary zone and the nozzle, and

(c) detonating the last named mixture in the secondary zone.

3, Apparatus for combusting bi-propellants, one of which is a liquid, in a rocket motor having a primary non-stoichiometric reaction zone and a spaced secondary stoichiometric reation zone downstream thereof, comprising:

(a) a single circular orifice disposed at the forward end of said motor adapted to inject said liquid in a direction axially thereof and parallel to the flow of combuston gases in the motor in the presence of the other propellant and in a coherent stream and at a velocity such that a portion of the stream is removed thereupon for reaction in the primary zone,

(b) a portion of said stream adapted to pass through said primary zone without reacting and to subsequently mix with the products of combustion from the primary zone to form a stoichometric mixture adapted to react in said secondary zone,

(c) said motor being provided with a convergentdivergent exhaust nozzle, the divergent portion having two divergent portions joined by a transition portion adapted to produce a Shock wave, said secondary zone being disposed within said nozzle rearwardly of said shock wave.

References Cited by the Examiner UNITED STATES PATENTS 2,996,880 8/1961 Greiner 60--35.6 3,077,073 2/1963 Kuhrt 60-35.6 3,093,960 6/1963 Tyson 60 35.4 3,164,093 1/1965 Holzman et al. 60-35.6

MARK NEWMAN, Primary Examner.

SAMUEL FEINBERG, Exam'ner.

Claims (1)

1. 3. APPARATUS FOR COMBUSTING BI-PROPELLANTS, ONE OF WHICH IS A LIQUID, IN A ROCKET MOTOR HAVING A PRIMARY NON-STOICHIOMETRIC REACTION ZONE AND A SPACED SECONDARY STOICHIOMETRIC REATION ZONE DOWNSTREAM THEREOF, COMPRISING: (A) A SINGLE CIRCULAR ORIFICE DISPOSED AT THE FORWARD END OF SAID MOTOR ADAPTED TO INJECT SAID LIQUID IN A DIRECTION AXIALLY THEREOF AND PARALLEL TO THE FLOW OF COMBUSTION GASES IN THE MOTOR IN THE PRESENCE OF THE OTHER PROPELLANT AND IN A COHERENT STREAM AND AT A VELOCITY SUCH THAT A PORTION OF THE STREAM IS REMOVED THEREUPON FOR REACTING IN THE PRIMARY ZONE, (B) A PORTION OF SAID STREAM ADAPTED TO PASS THROUGH SAID PRIMARY ZONE WITHOUT REACTING AND TO SUBSEQUENTLY MIX WITH THE PRODUCTS OF COMBUSTION FROM THE PRIMARY ZONE TO FORM A STOICHIOMETRIC MIXTURE ADAPTED TO REACT IN SAID SECONDARY ZONE, (C) SAID MOTOR BEING PROVIDED WITH A CONVERGENTDIVERGENT EXHAUST NOZZLE, THE DIVERGENT PORTION HAVING TWO DIVERGENT PORTIONS JOINED BY A TRANSITION PORTION ADAPTED TO PRODUCE A SHOCK WAVE, SAID SECONDARY ZONE BEING DISPOSED WITHIN SAID NOZZLE REARWARDLY OF SAID SHOCK WAVE.

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