US2972225A - Motor mechanism for missiles - Google Patents

Motor mechanism for missiles Download PDF

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US2972225A
US2972225A US199118A US19911850A US2972225A US 2972225 A US2972225 A US 2972225A US 199118 A US199118 A US 199118A US 19911850 A US19911850 A US 19911850A US 2972225 A US2972225 A US 2972225A
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tank
grain
chamber
rocket
aft
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US199118A
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James M Cumming
Gail M Dyer
Wolfgang C Noeggerath
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James M Cumming
Gail M Dyer
Wolfgang C Noeggerath
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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
    • F02K9/44Feeding propellants
    • F02K9/52Injectors
    • 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
    • F02K9/44Feeding propellants
    • 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
    • F02K9/95Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements

Description

Feb. 21, 1961 J. M. CUMMING ETAL 2,972,225
MOTOR MECHANISM FOR MISSILES Filed Dec. 4, 1950 3 Sheets-Sheet 1 dz, 1 a {@Wfi INVENTORS.
A T TORNEYS 3 Sheets-Sheet 2 Feb. 21, 1961 J. M. CUMMING ETAL MOTOR MECHANISM FOR MISSILES Filed Dec. 4, 1950 JNVENTORS. James M. Cumming Gail M. Dyer Wolfgang C. Naeggerafh UM. T Affomeys QM 7 7// A 1961 J. M. CUMMING EIAL MOTOR MECIJANISM FOR MISSILES Filed Dec. 4, 1950 3 Sheets-Sheet 3 K INVENTORS. James M. Cumming Gail M Dyer Wolfgang C. No eggerafh United States Patent MOTOR IVIECHANISM FOR MISSILES James M. Cumming, San Marino, Gail M. Dyer, China Lake, and WolfgangC. Noeggerath, El Monte, Calif.
Filed Dec. 4, 1950, Ser. No. 199,118
Claims. (Cl. fill-35.6)
(Granted under Title 35, US. Code (1952), sec. 266) 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 therefore.
This invention relates to motor mechanism for missiles, and is specifically directed although not necessarily limited to rockets.
It is one object of this invention to improvethe depend ability of engine combustion.
It is another object of this invention to improve the ignition of liquid propellants in a jet engine.
Another object of this invention is the reduction in the size of the Combustion chamber, eflected by better preparation of the propellants for combustion.
It is a further object of the instant invention to provide a jet engine capable of employing a high mass flow rate whereby a small combustion chamber may be employed.
In the past, the ignition of liquid propellants has involved either the employment of spontaneously reacting propellants, which are critical in their handling and operation, or the use of non-spontaneously reacting liquid propellants supplemented by use of catalysts, squibs, hot wires, or small pilot flames, all such igniters being designed to create small ignition centers which in turn raise the energy level in the combustion chamber to the point of full burning. It is an object of this invention to furnish initial energy in a jet motor in very large quantities and to prepare the liquid fuel principally by atomization thereof, so that full burning in the combustion chamber can be reached more rapidly than by previous methods. The use of very large quantities of energy for initiating combustion and the preparation of propellants for combustion--both of which are effected by this inventionare equally important in applications where motors burning spontaneously reacting propellants must attain full thrust immediately upon ignition.
Another object of the instant invention is to achieve a process of applying relatively high amounts of energy in a combustion chamber in an extremely short time to bring about combustion of liquid propellants. It is a concept of the instant invention that this ignition energy, in the form of an ignition flame, should not merely be used at the start of the injection of the liquid propellant, but should be used continuously during the burning of the liquid propellant so that the liquid always has available an ignition flame during its entire injection cycle in the burning chamber.
It is a further object of this invention to obviate the necessity for a separate pressurizing source for the liquid propellant. In the instant invention this is achieved by using the auxiliary or ignition propellant to also pressurize the tanks containing the principal liquid propellant.
It is an object of this invention to provide a rocket which employs liquid propellant as the primary source of power, and also provides a solid propellant to serve for initial thrust. The solid propellant not only provides initial thrust'for the jet-driven vehicle proper, but through r, 2,972,225 Patented Feb. 21, 1961 this initial thrust also serves to aid in feeding liquid propellant into the combustion chamber.
It is a further object of this invention to provide a fluid mixing chamber of improved design especially adaptable for use in a jet engine.
In the pressurizing of liquid propellants by hot gases from another burning propellant, it may under certain circumstances be advisable to isolate the hot gases from the propellant. To that end it is an object of this invention to provide isolating means which will prevent the hot gases from coming in contact with the fluid propellant, and at the same time will transmit the pressure of the hot gases to the fluid propellant to pressurize the fluid propellant tank.
Many of the teachings to be disclosed hereinafter are applicable to all types of vehicles propelled by the ejection of a stream of fluid. It is therefore to be understood that by the term jet engine is meant any propulsive means wherein force is generated by the reaction of a fluid jet issuing from the engine. While many of the concepts of the instant invention are uniquely adapted to rocket propulsion, it will be readily evident that the field of the instant invention is in jet engines generally, and not limited solely to rockets.
In employing fluid fuels, it is common practice to combine the fluid fuel with a fluid oxidant, the'chemical reaction between the two amounting to a burning of the fiuid fuel. This burning or combustion creates the jet which drives the rocket or other vehicle. Since, mechanically, accommodation of the fluid fuel and of the fluid oxidant involves generally the same problems, it is to be understood that by the term fluid propellant used herein is meant any fluid reactant used in the jet propelling of a jet engine, and propellant alone includes any and all reactants used in the engine.
Other objects and many of the attendant advantages or" this invention will be readily appreciated as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings, wherein:
Fig. 1 is an over-all view of a rocket, embodying several features of the instant invention;
Fig. 2 is an enlarged fragmentary section illustrating the tail portion of the rocket illustrated in Fig. 1;
Fig. 3 is a fragmentary sectional view showing the forward section of the rocket immediately abaft the cargo space or warhead;
Fig. 4 illustrates a toroidal water bag which may be used to isolate the oxidant from the hot pressurizing gases.
Although the invention is applicable to the ram jet type of engine, it will be herein described as applied to a 'fuel miXing-and-ignition chamber communicating:
forward or cargo portion 46, a pressurizing portion 47, a fuel supply portion 48, and a combustion chamber and nozzle portion 49. Primary features of the instant invention are to be found in the combustion chamber and nozzle portion 49; important features are to be found in the pressurizing portion 47, and in the fuel supply portion 48; and other important features are embodied in the combination of the three portions 47, 48, and 49, as will be more particularly pointed out hereinafter.
Referring to Fig. 2, a preferred embodiment of the combustion chamber and nozzle or aft portion 49 of the rocket of Fig. 1 is shown. In Fig. 2, 51 designates a with an elongate hollow powder grain 52 (only the extreme aft end of which is shown in Fig. 2) thru the intermediacy of a nozzle 53, (2) with an annular oxidant tank 54 longitudinally substantially coextensive with the powder grain 52 and disposed circumjacent the grain 52, and (3) with an annular fuel tank 56 likewise of substantially the same length as the powder grain 52, and surrounding the oxidant tank 54.
The forward end of the Housing 57 defining the chamber 51 has a conical-portion SSringing the mouth of the nozzle 53. Communication between the annular oxidant tank54 and thechamber 51 is effected through a series of first inlet ports 59, located in the conical portion 58 of the housing 57. The ports 59are formed so as to introduce oxidant from the tank 54 into the chamber 51 with a circumferentialcomponent, in order to impart wln'rling :motion' to the oxidant as it is ejected into the chamber 51. In Fig. 2 this whirlingmotion is effected by the provision of a plurality of defiectingvanes 61' adjacent the ports 59, which are slanted circumferentially, thereby imparting the desired whirling motion to the oxidant as it enters the chamber 51. I a r interposed between the end of the tank 54 and the conical portion 58 of the housing 57 is an annular wall 62 having a plurality of holes 63 therethru spaced on a circumference around the rocket axis. The annular wall 62 constitutes an aft end wall for the oxidant tank 54, which is formed by a pair of concentric aluminum tubes 64 and 66, sealed at the aft end by a thin annular sheet of aluminum 67, the internal and external edges of which are formed into locked seams 68 and 69 with the edges of the cylinders 66 and 64, respectively. The annulus 67, which forms a frangible wall to keep oxidant within the tank 54, is further secured to'the end edges of the tank 54 by a ring 71 having a plurality of ports 72 registering with the holes 63 in the end wall 62. The edges of the ports 72 are made sharp to aid in cutting the frangible wall 67 when the oxidant in tank 54 is pressurized. The ring 71 is pressed into place adjacent the outer lock seam 69 and the inner lock seam 68, being in turn held in place by the wall member 62.
Communication between the fluid fuel'tank 56 and the chamber 51 is provided by a series of second inlet ports 73, which are likewise formed to introduce fluid fuel to the chamber 51 with a circumferential component, whereby the fuel-like the oxidant-is made to whirl about the nature or" either the ports 59 or the ports 73. Like the ports 59, the ports 73 impart whirling motion to the fluid by virtue of deflecting vanes 74 slanted circumferentially adjacent the ports 73.
Annularly banded about the ports 73, so as to be interposed in the communication between the chamber 51 and the tank 56, is a band of plastic or metal foil 76, which forms a frangible wall keeping fluid fuel Within the tank 56 until such time as it should be ejected into the chamber 51, at which point a highpressure is built up in the tank 56 to fracture the band 76 opposite each port 73, and allow fluid fuel to be ejected into the chamber 51 through the ports 73.
The housing 57, tanks 54 and 56, and elongate powder grain 52 are held within the elongate cylindrical rocket shell 77 by means of an annular securing member 78 keyed to the shell 77 by a snap ring 80. Fluid tight sealing at the edges of theannular member 78 is provided by an external O-ring 79, and an internal O-ring 81.
It will be. seen from Fig. 2 that the annular fluid fuel tank 56 is in effect nothing more than the annular space between the exterior cylindrical wall 64 of the oxidant tank, and the shell 77 of the rocket proper.
Directly aft of the ignition-and-mixing chamber 51 is the combustion chamber 82, formed by lining that portion of the rocket shell 77 with a graphite liner 83. At the open aft end of the combustion chamber 82 is an exhaust nozzle 84, abutted against the aft edge of the graphite liner 83 by a screw ring 86 at the extreme end of the rocket.
The extremely high pressures generated within jet propulsion combustion chambers and nozzles has dictated that the walls of all parts be substantially rigid. This requirement has extended even to the relatively thick restrictive wall constituting the nozzle proper. It is a feature of the instant invention that the nozzle portion 84 of the rocket is made hollow, i.e., with an annular chamber 37, thereby ridding the rocket. of. some excess weight. In order to. prevent the walls of the nozzle 84 from collapsing outwardly into thehollow portion 87, there is provided a plurality of holes 88 passing through the wall of the nozzle 84 and communicating between the combustion chamber 82 and the interior 87 of the nozzle 84. The holes 88 serve as pressure-equalizing holes by means of which the static pressure in the interior 87 of the bosslike portion of the nozzle 84 may be made equal to the static pressure in the chamber 82. There is thus no tendency for the walls of the boss or nozzle to collapse.
The aft end of the rocket is provided with the usual stabilizing vanes 89 exteriorly thereof.
Turning now to Fig. 3, the portion 47 of the rocket called herein the pressurizing portion will now be described. It is to be understood that the propellant portion 48 of the rocket is similar in cross section throughout its'length, and consists simply of the powder grain 52, the circumjacent oxidant tank 54, and the circumjacent fuel tank 56, all extending the full length of the portion 48 between the nozzle portion 49 and the pressurizing portion 47.
The oxidant tank 54 is closed at its forward end by an annular end plate 91 having a plurality of holes 92 spaced circumferentially therearound. Covering the holes 92 is an annular frangible wall or burst diaphragm 93, held in place by pressing the ring or annular wall into the an nular opening at the forward end of the tank 54. The forward edges of the cylinders 64 and 66 forming the tank 54 are contained between two rings 94 and 96. A ring 99 screwed internally into the ring 94- and a ring 98 screwed externally onto the ring 96 serve to hold the annular wall 91 in place against the annular burst diaphragm 93 and against the tank walls 64 and 66. This assembly, coupled with the aft assembly 67, creates an individually sealed tank 54 whichmay be removed as a unit from the rocket. In this manner, the oxidant tank 54 may be assembled, filled, and sealed outside of the rocket, and need not be placed in position until just prior to the launching of the rocket. This feature not only adds to the safety of handling, but also precludes unnecessary corrosion due to long storage of the oxidant inside the tank 54.
An end closure ill-1 serves to complete a pressurizing communication between the powder grain 52, the tank 54, and the tank 56, to be more fully explained hereinafter. The closure 191 has formed integrally therewith an after-extending cylindrical flange H2 immediately circumjacent the ring 94 which forms part of the integral assembly of the tank 5 The joint 103 between the flange 1G2 and the ring 94 is made sliding, so that unequal longitudinal expansion of the central portion of the rocket relative to the exterior thereof may be accommodated. That is to say, the aft end of the central portion of the rocket including the powder grain 52 and the tank 54 is fixedly mounted with respect to the rocket shell 77. At the forward end, however, the mounting is made sliding on the surface 103, and sealed by an O-ring 100, so that relative longitudinal movement can take place between the'assembly 52-54 and the remainder of the rocket, including the shell 77 That portion of the flange v192 which is forward of the ring 94 is provided with anumber of ports or holes 104, so as to provide communication'between the tank 56 and the space 166' immediately aft of the closure Wall 103. The ports 104 are closed by a frangible wall in the form of a cylindrical strip of plastic foil 195, wound around the flange 102 immediately circumjacent the ports 164;
The powder grain 52 is hollow throughout its length so as to communicate not only with the mixing-and-ignition chamber 51 (Fig. 2), but also with the space 106 immediately aft of the closure wall 101. This communication is preferably effected through the intermediacy of a nozzle restriction 107, similar to the nozzle restriction 53 at the aft end of the grain 52.
It will thus be seen that the powder grain 52 communicates at its forward end with both the tank 54 and the tank 56, both communications being blocked by the interposition of frangible walls or burst diaphragms 93 and 105, respectively. The purpose of the communicating region 106 is to allow the grain 52 to pressurize the tanks 54 and 56. Thus when the grain starts to burn, a flame is ejected forwardly through the nozzle 107 creating a pressure in the space 106, shattering the frangible walls 93 and 105, and applying pressure to the fluid inside the tanks 54 and 56. This pressure in turn shatters the frangible walls 67 and 76 at the aft end of the tank causing the oxidant and fuel from the tanks 54 and 56 to be ejected into the chamber 51 in a manner to be more fully explained hereinafter.
Hot gases from the burning powder grain 52 shooting rapidly outward in the space 106 and impinging on the rocket shell 77 tend to burn or corrode the interior of the shell. To protect the shell at this point, it is lined with a refractory liner 108.
Ignition of the grain 52 is effected through a suitable igniter 109 resting interiorly of the grain 52, and energized by electrical leads 111 extending through the nozzle 107 and through the forward closure 101. It will be understood that application of an electric potential between the terminal 112 and the shell 77 (one of the leads 111 being grounded to the shell) will energize the igniter 109 to ignite the powder grain 52.
The tubular or hollow powder grain 52 is contained within a thin, elongate cylinder 113. Under the extremely high accelerative forces encountered during launching of the rocket, there is danger that the burning grain 52 might collapse and be driven to the rear where the fragments thereof would clog up the nozzle 53. To strengthen the powder grain against such forces, a plurality of annular strengthening partitions 114, of annular width equal to the thickness of the grain wall 52, are secured as by welding at 116 to the cylinder 113 and are spaced longitudinally therealong.
it will be noted that in the embodiment under discussion, the oxidant tank 54 is located within the fuel tank 56, rather than vice versa. because it is preferred to have the irlet ports 59 for the oxidant located upstream of the inlet ports 73 of the fuel. This disposition of inlet ports is to be preferred because the density of the oxidant is usually higher than that of the fuel, and therefore the swirling motion will tend to drive the heavier oxidant centrifugally outward into the lighter fuel whereby a better mixing of the two propellants is achieved. Interior disposition of the annular oxidant tank 54 can rest around and be supported by the relatively strong grain tube 113. Strengthening of the oxidant tank wall is advantageous as a protection against collapsing of the inner tank wall in case of an increase of pressure due to decomposition of the oxidant, such as fuming nitric acid, under prolonged storage conditions.
Operation The operation of the rocket will now be described with reference to Figs. 2 and 3. Before launching of the rocket, the oxidant filled tank 54 is placed within the rocket shell 77 and secured between end members 73 and 101 and their associated assemblies. The outer annular tank 56 is then filled with fuel. The rocket is then placed on a suitable launcher.
Application of electric potential to the leads 111 energizes the igniter 109 which initiates interior burning along the length of the powder grain 52. Burning of the grain This is desirable principally ejects a flame rearwardly through the nozzle 53, which continues out the principal nozzle 84, thereby giving thrust to the rocket and starting it on its way.
Burning of the grain 52 also ejects a flame forwardly into the space 106, thereby building up a high pressure therein. This pressure serves to fracture the frangible walls or diaphragms 93 and 105, thereby placing the oxidant in the tank 54 and the fuel in the tank 56 under pressure from the forward end. This pressure bursts the frangible walls 67 and 76 at the aft end of the two tanks and ejects oxidant and fuel into the mixing-andignition chamber 51. Fuel ejection is further heightened by the acceleration of the rocket itself (due to the thrust flame from the nozzle 53), because under acceleration the contents of the tanks 54 and 56 are driven to the after end of the tank.
The nozzle 53 thus serves two very important functions. It gives initial thrust to the rocket during the time taken to mix and ignite the liquid fuels in the chamber 51, and it also serves to create a higher pressure in the interior of the powder grain 52 than in the chamber 51. This higher pressure is needed in the pressurizing of the tanks 54 and 56, for otherwise there would be no pressure differential to drive the propellant from the tanks 54 and 56 into the chamber 51. V
The initial thrust afforded by the nozzle 53 is also important to assure that the pressurizing gases from the space 106 (now in communication with the tanks 54 and 56 through the bursting of the diaphragms 93 and do not pass rearwardly through the liquid in the tanks 54 and 56. Thisis assured by providing that the net, or aggregate, acceleration of the rocket is always such as to drive the contents of the tanks to the aft end of the rocket. For example, were the rocket to be fired upwardly from a stationary position, there would be no problem encountered, for the contents of the tanks 54 and 56 would normally be pulled by gravity to the aft end of the rocket, and the pressurizing gases would have no tendency to pass through the tanks. if, however, the rocket were to be pointed down upon firing, then it is most essential that the powder grain 52 impart to the rocket an acceleration of at least one g, for otherwise the liquid propellants in the tanks 54 and 56 would tend to go to the forward or bow end of the rocket, and the pressurizing gases would bubble through the liquid to the aft end.
Propellants ejected from the tanks 54 and 56 into the chamber 51 are given a whirling motion by virtue of the deflecting vanes 61 and 74. Thus a spray of liquid oxidant from the tank 54 and of liquid fuel from the tank 56 is mixed and whirled about the chamber 51, the whirling motions in this case being in the same direction. The axial flame jetted from the nozzle 53 has two functions with respect to the swirling spray from the ports 59 and 73. Firstly, it tends to atomize the liquid propellants into an extremely fine spray; and secondly, it raises the temperature of the propellants to their reaction or ignition point, so that they combine and produce the principal combustion tending to propel the rocket.
As seen in Fig. 2, the aft end of the housing 57 is curved inwardly at 121, so that the swirling spray, as it moves rearwardly to the mouth of the chamber 51, is deflected radially inward into the flame from the nozzle 53 where complete and even ignition is achieved. The open mouth 121 of the housing 57 is provided with a plurality of longitudinal bafiles 122 secured interiorly thereof and spaced circumferentially therearound. It is the function of the bafiles 122 to transmute at least a portion of the swirling motion of the propellants into a turbulence which creates better mixing of the oxidant and fuel.
The greater portion of the actual burning, or chemical reaction, between the liquid oxidant and fuel takes place in the combustion chamber 82, the principal function of the chamber 51 being to mix and ignitethe two propellants. Full combustion thus takes place in the chamber that the flame will be extinguished before complete utili-' zation of the liquid propellants, such as might occur were self-combustion to be relied upon after initial ignition of the two. liquids.
Mixingof the propellants in the chamber 51 primarily in the liquid phase rather than in the vapor phase is to be desired because of the fact that, in restricted space, liquid mixing is generally more effective and results in a more homogeneous end product than does vapor mixing.
The nozzle restriction 107, while improving the burning characteristics of the powder grain 52, is not essential for satisfactory operation of the rocket illustrated. During the burning of the grain 52, the interior portion of the rocket achieves a temperature considerably higher than the exterior. The resulting unequal expansion of parts is accommodated, as hereinbefore explained, by the sliding mounting at the forward end of the tank 54, where the ring 94 is mounted for longitudinal movement within the flange 102 of the closure plate 101.
In case the composition of the powder grain 52 is such that it has a negative oxygen balance, it is highly desirable and in fact perhaps necessary, to isolate the hot gases in the space 106 (Fig. 3) from the oxidant in the tank 54. A method of achieving such isolation is illustrated in Fig. 4 wherein there is positioned in the forward end of the tank 54' a toroidal bag 131 made for example at plastic and filled with water or other inert liquid having a lower density than the oxidant. The bag 131 is thus interposed between the oxidant in the tank 54' and the hot grain gases in the space 106. In actual operation the bag 131 rapidly shatters, but this in noway destroys the eflicacy of the isolating water layer which, being less dense than the oxidant, tends to float on the surface thereof. It will be understood that this floating action takes place even though the rocket may be actually traveling downward, because the downward acceleration of the rocket is greater than one g. The annular column of water thus serves to efiectively isolate the pressurizing gas from the oxidant, and serves as apiston for transmitting pressure between the two fluids without allowing them to come into contact, with consequent danger of chemical reaction and combustion.
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 scope of the appended claims the invention may be practiced otherwise than as specifically described.
What is claimed is:
1. In a jet engine, the combination of an elongate hollow powder grain, a first annular fluid propellant tank circumjacent said grain, the forward end of said grain communicating with the forward end of said tank, a frangible wall interposed across the communication between said grain and said tank, thereby to contain the propellant within the tank, a housing defining an ignition chamber disposed aft of said grain and communicating therewith, restriction means interposed across the communication between said grain and chamber, said chamber being open at its aft end for the ejection of mixed propellants, the forward end of said housing having a conical portion centered about a central axis passing through said open end, a series of first inlet ports in said conical portion spaced about the chamber annularly with respect to said central axis, a series of second inlet ports spaced about the chamber annularly with respect to said axis and displaced longitudinally toward said open end from said series of first ports; deflecting vanes adjacent said inlet ports slanted circumferentially to impart whirling motion to fluids entering the chamber through said ports,'the aft end of said tank communicating with said first inlet ports for injecting first propellant into said chamber, a frangible Wall interposed across the communication between said chamber and said aft end thereby to contain the propellant within the tank, a second annular propellant tank circumjacent said first propellant tank, the forward'end of said second tank also communicating with the forward end of said grain, the aft end of said second tank communicating with said second inlet ports for injecting second propellant into said chamber, means for igniting said grain whereby to initiate an ignition flame from said grain substantially axially into said chamber, combustion ofsaid grain simultaneously shattering said frangible'walls thereby to eject propellant from the aft ends of said tank into said chamber, the open end of said housing converging toward said axis thereby to direct mixed propellant inwardly toward the ignition flame, and a plurality of baflles secured to the interior of said housing and spaced circumferentially therearound near the open end thereof effective to transmute at least some of the whirling motion of the fluids into turbulence for better mixing of the two propellants.
2. In a jet engine, the combination of an elongate powder grain, a first annular fluid propellant tank circumjacent said grain, the forward end of said grain communicating with the forward end of said tank, a housing defining an ignition chamber disposed aft of said grain and communicating therewith, restriction means interposed across the communication between said grain and chamber, said chamber being open at its aft end for the ejection of mixed and ignited propellants, a series of first inlet ports spaced about the chamber, a series of second inlet ports spaced about the chamber, the aft end of said first tank communicating with said first inlet ports for injecting first propellant 'into said chamber, a second annular propellant tank circumjacent said first propellant, tank, the forward end of said second tank also communicating with the forward end of'said grain, the aft end of said second tank communicating with said second inlet ports for injecting second propellant into said chamber, and means for igniting said grain whereby to initiate an ignition flame from said grain substantially axially into said chamber, combustion of said grain simultaneously pressurizing the forward ends of said tanks to eject propellant from the aft ends thereof into said chamber.
3. In a jet engine, the combination of an elongate cylindrical shell, a cylindrical powder grain disposed axially within said shell, an elongate annular liquid tank dis posed co-axially circumjacent said grain, a pressurizing housing disposed forward of said grain and communicating with the forward end of said grain and said tank, whereby burning of said grain serves to pressurize the liquid in the tank at the forward end of the tank, thereby to eject liquid from the aft end of said tank, and an inert liquid filled toroidal bag disposed in said-tank near the forward end thereof and filling the cross-sectional area of the tank, thereby to serve as a piston for transmitting pressure from the burning grain to the tank liquid, and to isolate said liquid from the pressurizing gases.
4. In a jet engine, the combination of an elongate cylindrical shell, a cylindrical hollow powder grain disposed axially within said shell, a pair of elongated annular liquid fuel tanks disposed co-axially and concentric to said grain,
an ignition housing disposed aft of said grain and communicating with the aft end of said grain and said tanks, restriction means interposed across the communication between said grain and housing, a pressurizing housing disposed forward of said grain and communicating with the forward end of said grain and said tanks, whereby burning of said grain serves to pressurize the liquid fuel in said tanks at the forward ends of said tank, thereby to eject liquid fuel into said ignition housing where it is ignited by the burning of said grain from its aft end, diaphragm means interposed in the communication between said tanks and said ignition housing effective to burst upon application of pressure from the liquid fuel, and diaphragm means interposed in the communication between said tanks and said pressurizing housing effective to burst upon application of pressure from the forward end of the hollow burning grain.
5. In a jet engine, the combination of an elongate cylindrical shell, a cylindrical hollow powder grain disposed axially within said shell, an elongate annular liquid propellant tank disposed co-axially and concentric to said grain, an ignition housing disposed aft of said grain and communicating with the aft end of said grain and said tank, restriction means interposed across the communication between said grain 'and housing, and a pressurizing housing disposed forward of said grain and communicating with the forward end of said grain and said tank, whereby burning of said grain serves to pressurize the liquid propellant in the tank at the forward end of the tank, thereby to eject the liquid propellant into said ignition housing where it is ignited by the burning of said grain from its aft end.
6. In a jet engine, the combination of an elongate hollow powder grain, an annular fluid propellant tank circumjacent said grain, the forward end of said grain communicating with the forward end of said tank, a restriction in the form of a first nozzle interposed in the communication between the forward end of said grain and the forward end of said tank, a frangible wall interposed across the communication between said grain and said forward end, thereby to contain the propellant within the tank, a housing defining an ignition chamber disposed aft of said grain and communicating therewith, a restriction in the form of a second nozzle interposed in the communication between the aft end of said grain and said chamber, the aft end of said tank communicating with said housing whereby propellant may be injected into said chamber, a frangible wall interposed across the communication between said chamber and said aft end, thereby to contain the propellant within the tank, and means for igniting said grain whereby toinitiate an ignition flame from said grain through said nozzle into said chamber, the combustion of said grain simultaneously shattering said first named frangible wall, thereby to pressurize the forward end of said tank and eject propellant from the aft end thereof into said chamber.
7. In a jet engine, the combination of a double walled, generally cylindrical tank containing fluid in .the annular interspace between the walls, a source of pressurizing fluid communicating with one end of said interspace to force the fluid therein out the other end of the tank, and an inert liquid filled toroidal bag interposed between the two mentioned fluids and filling the cross-sectional area of the interspace, thereby to serve as a piston for transmitting pressure from the pressurizing fluid to the tank fluid.
8. In a jet engine, the combination of a generally cylindrical tank containing fluid, a source of pressurizing fluid communicating with one end of said tank to force the fluid therein out the other end of the tank, and an inert liquid filled bag interposed between the two mentioned fluids and filling the cross-sectional area of the tank, thereby to serve as a piston for transmitting pressure from the pressurizing fluid to the tank fluid.
9. A rocket motor comprising an elongated powder grain having a longitudinal bore extending in fore-and-aft direction therethrough, a fluid propellant tank located adjacent the powder grain in generally parallel relation and concentric thereto, a passage connecting the forward ends of said bore and said tank, means defining an ignition chamber aft of said bore and said tank and in communication with the aft ends of said bore and said tank, and means for igniting the surface of said bore thereby to project an ignition flame from the aft end of said bore into said chamber and from the forward end of said bore to apply pressure to the forward end of said tank to thereby force fluid propellant into said chamber to be ignited by said flame.
10. A rocket comprising a first elongated tube having its longitudinal axis disposed on the longitudinal axis of the rocket and providing a first space therewithin, a second elongated tube concentrically surrounding the first tube and providing a second and annular space disposed around the first tube adapted to contain a liquid oxidant, a third elongated tube concentrically surrounding the second tube and providing a third and annular space disposed around the second tube adapted to contain a liquid fuel, wall means closing the front end of said second and third spaces, a chamber disposed adjacent the front end of the first tube and communicating therewith, frangible means associated with said wall means adapted to be ruptured by pressure and establish communication be tween said chamber and said second and third spaces at the forward ends thereof, second wall means closing the rear end of said second and third spaces, an axially aligned mixing and ignition chamber disposed adjacent the rear end of the tubes and communicating with the first tube, frangible means associated with said second wall means adapted to be ruptured by pressure and establish cornmunication between said mixing and ignition chamber and the rear ends of said second and third spaces, an axially aligned combustion chamber disposed rearwardly of said mixing and ignition chamber, an elongated powder grain disposed within said first space having an axial bore extending between its opposite ends, the forward end communicating with the first named chamber and the rearward end communicating with said mixing and ignition chamber, said grain being constructed to burn for the duration of operation of the rocket and provide thrust for same, provide a continuing ignition flame exhausting axially through the mixing and ignition chamber for maintaining ignition, and for providing continued pressure to the front ends of said second and third spaces whereby the liquid oxidant and propellant are continuously fed to said mixing and ignition chamber, said mixing chamber including means for injecting the liquid fuel and oxidizer in swirling directions around its longitudinal axis to thereby facilitate initial mixing and ignition, and angularly spaced bafiie means within the mixing and ignition chamber disposed rearwardly of the injecting means, for transforming at least some of the swirling motion into turbulence for further intimate mixing and final combustion in the combustion chamber.
References Cited in the file of this patent UNITED STATES PATENTS 47,544 Hotchkiss May 2, 1865 622,479 Isham Apr. 4, 1899 807,494 Du Pont Dec. 19, 1905 1,436,018 Dieter Nov. 21, 1922 1,692,710 Spahn Nov. 20, 1928 1,877,983 Schilling Sept. 20, 1932 1,879,186 Goddard Sept. 27, 1932 1,960,810 Gordon May 29, 1934 1,991,390 Holzwarth Feb. 19, 1935 2,090,039 Goddard Aug. 17, 1937 2,342,096 Zimmerman Feb. 15, 1944 2,345,540 Ray Mar. 28,1944 2,398,125 Summerfield et a1. Apr. 9, 1946 2,402,826 Lubbock June 25, 1946 2,408,111 Truax et al Sept. 24, 1946 2,419,866 Wilson Apr..29, 1947 2,447,758 Lubbock et al Aug. 24, 1948 2,478,958 Wheeler et a1 Aug. 16, 1949 2,481,059 Africano Sept. 6, 1949 (Other references on following page) 11 UNITED STATES PATENTS V Goddard Dec. 27, 1949 Goddard Feb. 21, 1950 Skinner May 2, 1950 Lauritsen July 11, 1950 5 Zucrow Aug. 29, 1950 Goddard Sept. 26, 1950 Goddard Oct; 17, 1950 Goddard Jan. 2, 1951 Simms Apr. 10, 1951 10 Nedoh May 15, 1951 Chilton 2 Feb. 12, 1952 Roy ....2 Mar. 9, 1954 FOREIGN PATENTS Ita1y Aug. 22, 1930 Great Britain July 11, 1921 OTHER REFERENCES pages 9-11.
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Cited By (17)

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US3094071A (en) * 1959-06-30 1963-06-18 Union Carbide Corp Vacuum insulated storage tanks for missile use
US3094837A (en) * 1957-02-19 1963-06-25 Thiokol Chemical Corp Rocket motor
US3132475A (en) * 1961-06-29 1964-05-12 United Aircraft Corp Hybrid rocket propulsion system
US3138001A (en) * 1962-10-11 1964-06-23 William I Berks Device to eliminate shear slide in prepackaged liquid powerplants
US3178885A (en) * 1961-06-12 1965-04-20 Lockheed Aircraft Corp Hybrid rocket engine
US3215365A (en) * 1963-04-30 1965-11-02 Wyatt Theodore Spacecraft propulsion concept
US3253408A (en) * 1963-09-23 1966-05-31 Hollas K Price Rocket engine fuel feeding system
US3340691A (en) * 1965-10-14 1967-09-12 Thiokol Chemical Corp Command controllable self-pressurizing liquid injection system
US3368354A (en) * 1963-12-18 1968-02-13 United Aircraft Corp Rocket motor
US3374623A (en) * 1965-11-16 1968-03-26 Army Usa Method of operating a liquid oxidizer feed piston
US3377801A (en) * 1964-11-18 1968-04-16 United Aircraft Corp Liquid propulsion system and method with fuels and oxidizer in thermal contact
US3468487A (en) * 1966-02-28 1969-09-23 Us Navy Variable thrust injector
US3517511A (en) * 1967-12-09 1970-06-30 Rolls Royce Bi-propellant rocket engine
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
US20080241781A1 (en) * 2005-10-28 2008-10-02 Sefmat Rue De Betnoms Hot Air Internal Ignition Burner/Generator
US20120233979A1 (en) * 2011-03-16 2012-09-20 Raytheon Company Rocket multi-nozzle grid assembly and methods for maintaining pressure and thrust profiles with the same
RU2561796C1 (en) * 2014-10-16 2015-09-10 Открытое акционерное общество "Конструкторское бюро химавтоматики" Liquid-propellant rocket engine (lpre) combustion chamber with electroplasma ignition

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Cited By (19)

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Publication number Priority date Publication date Assignee Title
US3094837A (en) * 1957-02-19 1963-06-25 Thiokol Chemical Corp Rocket motor
US3094071A (en) * 1959-06-30 1963-06-18 Union Carbide Corp Vacuum insulated storage tanks for missile use
US3178885A (en) * 1961-06-12 1965-04-20 Lockheed Aircraft Corp Hybrid rocket engine
US3132475A (en) * 1961-06-29 1964-05-12 United Aircraft Corp Hybrid rocket propulsion system
US3138001A (en) * 1962-10-11 1964-06-23 William I Berks Device to eliminate shear slide in prepackaged liquid powerplants
US3215365A (en) * 1963-04-30 1965-11-02 Wyatt Theodore Spacecraft propulsion concept
US3253408A (en) * 1963-09-23 1966-05-31 Hollas K Price Rocket engine fuel feeding system
US3368354A (en) * 1963-12-18 1968-02-13 United Aircraft Corp Rocket motor
US3377801A (en) * 1964-11-18 1968-04-16 United Aircraft Corp Liquid propulsion system and method with fuels and oxidizer in thermal contact
US3340691A (en) * 1965-10-14 1967-09-12 Thiokol Chemical Corp Command controllable self-pressurizing liquid injection system
US3374623A (en) * 1965-11-16 1968-03-26 Army Usa Method of operating a liquid oxidizer feed piston
US3468487A (en) * 1966-02-28 1969-09-23 Us Navy Variable thrust injector
US3517511A (en) * 1967-12-09 1970-06-30 Rolls Royce Bi-propellant rocket engine
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
US20080241781A1 (en) * 2005-10-28 2008-10-02 Sefmat Rue De Betnoms Hot Air Internal Ignition Burner/Generator
US8678816B2 (en) * 2005-10-28 2014-03-25 Sefmat Hot air internal ignition burner/generator
US20120233979A1 (en) * 2011-03-16 2012-09-20 Raytheon Company Rocket multi-nozzle grid assembly and methods for maintaining pressure and thrust profiles with the same
US8596040B2 (en) * 2011-03-16 2013-12-03 Raytheon Company Rocket multi-nozzle grid assembly and methods for maintaining pressure and thrust profiles with the same
RU2561796C1 (en) * 2014-10-16 2015-09-10 Открытое акционерное общество "Конструкторское бюро химавтоматики" Liquid-propellant rocket engine (lpre) combustion chamber with electroplasma ignition

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