US2724237A - Rocket projectile having discrete flight initiating and sustaining chambers - Google Patents

Rocket projectile having discrete flight initiating and sustaining chambers Download PDF

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US2724237A
US2724237A US652206A US65220646A US2724237A US 2724237 A US2724237 A US 2724237A US 652206 A US652206 A US 652206A US 65220646 A US65220646 A US 65220646A US 2724237 A US2724237 A US 2724237A
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rocket
propellent
burning
charge
initiating
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Clarence N Hickman
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Clarence N Hickman
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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
    • F02K9/74Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof combined with another jet-propulsion plant
    • F02K9/76Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof combined with another jet-propulsion plant with another rocket-engine plant; Multistage rocket-engine plants
    • F02K9/763Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof combined with another jet-propulsion plant with another rocket-engine plant; Multistage rocket-engine plants with solid propellant
    • 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
    • F02K9/08Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/28Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants having two or more propellant charges with the propulsion gases exhausting through a common nozzle
    • 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
    • F02K9/08Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32Constructional parts; Details
    • F02K9/36Propellant charge supports
    • 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
    • F02K9/08Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32Constructional parts; Details
    • F02K9/38Safety devices, e.g. to prevent accidental ignition

Description

Nov. 22, 1955 C, N HlCKMAN 2,724,237

ROCKE ROJECTILE HAVING DISCRETE FLIGHT 1N ATING AND SUSTAINING CHAMBERS Filed March 5, 1946 2 Sheets-Sheet l Ellnr En C1. E1 N. Hi @kn-Lun @www Nov. 22, 1955 cz. N. HlcKMAN ROCKET PROJECTILE HAVING DISCRETE FLIGHT INITIATING AND SUSTAINING CHAMBERS 2 Sheets-Sheet 2 Filed Maron 5, 1946 1 vuol 11o/L lnrencne N. Hi :k1-nun.

2,724,237 Patented. Nov. 22,., 1 955 iiice ROCKET PROJECTILE `HAY/'ING `DISCRETE FLIGHT IN lTIATING` AND SUSTAINING i CHAMBERS Clarence N. Hickman, Jackson Heights, N. Y., assigner to the United States `of` America asrepresenterl by the Secretary of War The invention described herein may be manufactured and used by or for the Government for governmental purposes, withoutthe payment to me ofany royalty thereon.

This invention relates to` rocketV propelled projectiles and more particularly to` rocket propelled projectiles which are adapted to be fired fromprojectors held in the hands of ground troops.`

One of the foremost obstacles to` be` overcome in perfecting this type of rocket propelled projectile is that of eliminating the presence of a blast from the jet, after the rocket has leftthe projector, without causing a material reduction in the rockets nal velocity. The presence of this jet blast is a serious source of annoyance to the tiring personnel who must be-stationed, during firing, at a position to the rear of the muzzle `of thejprojector and in the direction of the line of ightl of` the rocket. Attempts have been made to protect the firing personnel from the eiiects of the jetblastbyproviding a` discshaped shield at the muzzle of the projector. it was found, however, that excessive recoil forces were produced by the gases of combustion striking the shield.` Consequently, use of this shield as a means for protecting the tiring` personnel has been largely abandoned.

The blast whichl occurs after the projectile has left the tiring tube can be eliminated only if the burning of the propellent is completed while the rocket is within the tube. lf high velocity and long range are not required the annoyance caused by the jet blast may be easily avoided by selectinga powder of web thickness (distance between burning surfaces) suiiiciently thin to have the burning completed during the time it takes `the `rocket to travel the length` of the projector. It is when `the velocity and range are to be increased that difficulties arise, since in order to increase the velocity, and still` have the burning completed by the time the projector has left` the tube,

x the powder charge must be increased.` But, by increasing the .powder charge, the quantity of gas ejected per unit time is also increased, with the result thathigher pressures `will be devoloped .in the `motor chamber unless the throat area of the nozzle is increased. it must be noted also that the total time for burning the propellent is decreased with an increase in velocity, since it takes less time for the rocket to travel the length of the tube. Therefore, on` obtaining a `higher velocity by increasing the charge, the web thickness must be further reduced to insure that the burning will be` over by the time the rocket leaves the tube. This reduction in web thickness increases the burning area of the propellent` with the result that still more gas is liberated per unit time. To provide for the ejection of this increase in gas liberation, the throat area must be again enlarged, otherwise higher pressures will be developed. Obviously, there is a limit to the amount the throat may be increased, It' cannot be made larger than `the, chamber. In fact, for practical as well as theoretical reasons, the throat diameter should not be more than one-third to one-half the inside diameter of the combustion chamber` In actual. practice; there is a limit to` the amountithe web thickness maycbe decreased without encountering ditliculties in trapping the propellent.. In `contrast to the burning of a propellent in a gun, the propellent in a rocket combustion chamber is not entirely consumed, i. e., unburned particles escape from the trap after the web of the powder has been reducedto a certain thickness. The smaller the original web thickness, the larger is the per.- centage of unburned powder that escapes, consequently, decreasing the web thickness is uneconomical and inefficient.

There are certain advantages to be obtained by allowing the pressure to increase with increase in powder charge. The higher pressures cause an increase in burning rate and a reduction in total burning time. The foremost objection to a high gas pressure in rockets is that a heavy walled combustion chamber must be providedto withstand these pressures. Obviously, the weight ofthe chamber cuts down the efiiciency and the ultimate range of the rocket.

Accordingly, it is an object of this invention to provide a rocket projectile which can be fired from a shoulder gun projector or other such projector adapted to be held in the hands of ground troops without having `the initial jet blast continued after the projectile has left the projector. Furthermore, the velocity and range of the rocket projector of this invention are not limited by the burning time of the propellent utilized to start the rocket in Hight.

A particular object of this invention is to provide an improved military rocket having two :separate driving charges. The first driving charge is a relatively quick burning charge, which imparts to the rocket projectile a thrust suiticient to get the projectile out of the projector tube and started in the direction of its normal trajectory. After the rocket projectile has cleared the projector by a safe, precalculated distance, dependent upon the size of the rocket and its particular application, the second driving charge is set off. This` second charge need not be quick burning but must be set off a predetermined interval after the burning of the first charge has been completed.

The specific nature ofthe invention as well as other objects and advantages thereof will clearly appear from a description of a preferred embodiment as shown in the accompanying drawings in which:

Fig. l is a fragmentary longitudinal sectional view of a rocket projectile embodying this invention;

Fig. 2 is a side elevational view of the explosive head partly broken away to show the fuse;

Fig. 3 is a cross sectional view taken along the line 3-3 of Fig. 1;

Fig. 4 is a cross sectional view taken along the line 4-4 of Fig. 1;

Fig. 5 is a view similar to Fig. i showing a modified form of the invention; i

Fig. 6 is an enlarged detail view in longitudinal section showing the thermal igniter utilized to effect discharge of the propellent charge;

Fig. 7 is an enlarged detail view showing the unitary grid utilized to trap the secondary propellent charge;

Fig. 8 is a cross sectional View taken along the line 8 3 of Fig. 5; and

Fig. 9 is a fragmentary sectional View taken along line 9-9`of Fig. 4.

Referringnow to Fig. 1, the high explosive and armorpiercing head 1 of the rocket has secured thereto a cylindrical sleeve 2 which is provided with an undercut threaded portion 22 for holding the head 1 to the trap plate 7 of casing `5. The thermal arming device 15 for the impact fuze 20 is threaded into the cylindrical sleeve 2 at itsA forward end. A screw 3 which isprovided with'a tapped threaded axial opening 23 is threaded into the base of the cylindrical sleeve 2 so that the closed end of the screw 3 projects into the cylindrical casing 5 when the cylindrical sleeve 2 is threaded into the trap plate 7. The firing pin 6 is provided with a cylindrical rod 24 of a diameter less than the diameter of the tapped opening 23. The end of this rod is threaded and adapted to be inserted into the opening 23. A small slug 25 of a low melting point alloy, such as Woods metal, is placed in the opening 23 of screw 3 so that when the temperature of such screw is raised to the melting point of the metal 25 and the rod 24 is inserted into this opening, it will remain firmly in place after the metal 25 solidies on cooling.

The cylindrical casing has its forward end threaded onto the periphery of trap plate 7 and has the walls at its rear end swaged inwardly to form an inwardly ilared entrance to a constricted throat portion 16 and then the walls taper outwardly in a rearward direction to provide a flared nozzle 17 for expanding the gases of combustion liberated by the propellent charge 18. This propellent charge is Vmade up of a plurality of cylindrical grains 19 of double-base powder, each having an axial cylindrical perforation. The grains are supported in the chamber in any suitable manner, such as by means of a plurality of wormshead trap wires 8, similar to those described in my copending application, Serial No. 538,315 filed June l, 1944, now abandoned, which have their forward ends threaded into trap plate 7 and upon which the propellent charges are strung.

An annular adapter 26 is provided having a generally frustro conical shape. This adapter has a plurality of radially divergent rib members 38 (Fig. 4) formed on its lateral surface which also have outer surfaces of frustro conical shape to conform to the shape of the nozzle flare 17. It should be noted that while the outer surface of the radially divergent ribs conforms to the interior configuration of the nozzle, the slope of the conical sides of the portion of the adapter shown in section is less than the slope of the nozzle surface and thus provides an annular nozzle space 37 of increasing area in the direction of the exit gases. The forward portion of the adapter is reduced in diameter to form the shoulders 27 and 28 and the extreme forward end 29 is threaded to engage in a threaded axial hole 30 in a star-shaped guide 9 which centers and supports the adapter 26 within the nozzle of the combustion chamber.

The adapter 26 has a rearwardly opening, conical bore 75 communicating with an axial hole 39 therethrough. The rear end of hole 39 is threaded to receive a threaded delay type thermal igniter which is preferably con* st ructed in a manner similar to the thermal igniter disclosed in my copending application Serial No. 538,314, filed. lune 1, 1944, Patent No. 2,459,163, ianuary 1S, 1949. The igniter comprises a threaded element having a forwardly opening axial powder recess 33 therein which aligns with hole 39. The powder recess 33 and hole 39 are filled with powder. A black powder primer 4 is suitably mounted within casing 5, preferably opposite the forward open end of hole 39.

The base of the adapter 26 is interiorly threaded to receive a secondary cylindrical casing 12. The cylindrical casing 12 is threaded into the rear end of adapter 26. This casing contains a relatively quick-burning propellent charge 34 and has a venturi nozzle 35 threaded on to an end thereof and supporting the flight stabilizing ns 13. The nozzle 35 also holds the grid trap 36 in place between the end surface of the cylindrical casing 12 and the venturi nozzle 35. An electrically red squib igniter 11 is suitably mounted within secondary cylindrical casing 12 and wires 14 connect squib 11 to conventional external contacts.

It should be noted here that the rocket employs the principles of design described and claimed in my copending application, Serial No. 576,439, tiled February 6, .1945, Patent No. 2,503,271, April 11, 1950, inasmuch as the center of thrust, produced by the high velocity discharge of the gases of combustion through the expanding annular passage 37, is substantially coincident with the center of gravity of the rocket. If any nonuniformity occurs in the discharge of gases through the nozzle 35, no dispersion will result because the direction of its line of flight is restricted by the projector, since the burning of the propellent 34 within secondary casing 12 is completed within the projector tube.

The operation of the described rocket embodying this invention is as follows: the web thickness of the propellent 34 is selected so that, for the particular operating gas pressure within the casing 12, the burning will be completed before the rocket leaves its projector tube. The weight of propellent 34 is just sufficient to give the rocket an initial thrust to get the rocket out of the projector and on its way toward the target. The heat generated by the burning propellent 34 will cause the black powder in the thermal igniter 10 to be ignited by thermal conduction through the walls of the thermal element. The time required for ignition is about O.i sec. in the embodiment illustrated. Due to the thrust produced by the charge 34, the rocket will have traveled a distance of about 50 feet in such time interval. The igniter 10 thus sets off igniter 4 which in turn ignites the primary charge 18. The gases produced by cornbustion of primary charge 18 discharge through armular nozzle space 37 and accelerate the projectile sutiiL ciently to insure that it will reach the target. About 0.1 sec. after the ignition of charge 18, suicient heat is generated within casing 5 by the burning propellent to melt solder 25 contained within the opening 23 of the screw 3 to release the rod 24 and ring pin 6. On impact the pin 6 and its inertia element are free to move forward, striking and exploding the percussion primer to set off a powder train leading to the high explosive in the head. The burning of the propellent 18 may be selected to give the rocket suiiicient velocity for penetration of the target armor by the specially designed armor-piercing head 1. j

A modification of this invention is shown in Fig. 5. The rocket comprises an explosive head 70, a primary combustion chamber 42 defined by a hollow, cylindrical, forward body member 4l) and containing the primary propellent charge 43 strung upon the worm head trap wires 44. A partition member 45 is threadably secured between forward body member 40 and a similar rear body member 41 and seals the main combustion chamber 42 from the fluid pressure developed in the secondary combustion chamber 46 by combustion of the quick burning secondary propellent 47. The secondary propellent 47 is held in the secondary combustion chamber 46 by means of a grid trap 48, which is retained against a shoulder formed at the rear end of chamber 46 by the nozzle 49 which is threadably secured thereto. The grid 48 is conveniently formed from a circular plate by milling parallel grooves 60 (Fig. 7) on one side of the plate in one direction to form the lands 61 and by milling parallel grooves 62 on the other side of the plate, in a direction at right angles to the direction of milling for the grooves 60, to form the lands 63. The result is a unitary grid member formed of cross bars having their surfaces projecting on opposite sides of a center line.

The partition member 45 is provided with a plurality of circular openings 50 for permitting the combustion gases liberated by the propellent 43 to pass rearwardly 0f the chamber for exit through the nozzle 49. The openings 50 are of equal diameter and are equally positioned on equally spaced radii of the partition member 45. In order to prevent ignition of the propellent 43 by the gases liberated from the quick burning propellent 47 in the combustion chamber 46, individual plate members 52 are provided to cover each of the openings 50 .in Athe partition member 45. `These plates each have a tab 53 formed integrally therewith for receiving a bolt 54 which screws into partition member 45 to retain the plate in place over the opening on the side of the partition facing the nozzle.

A thermal ignition device 55 is threaded into an axial hole 57 in the partition member 45 so that its closed end projects into the secondary combustion chamber 46. An ignition train S6 is suitably supported within primary combustion chamber 42 extending forward along the axis of the rocket body and serves to ignite black powder igniter 58 which in turn ignites the main pro pellent 43. A conventional electrically red squib 100 surrounds thermal ignition device 55 to ignite such device as well as secondary charge 47. Lead wires 101 and 102 connect squib 100 to the source of current (not shown).

In operation, the propellent charge 47, which is of a selected thin web, is consumed by burning in a time interval less than that required for the projectile to travel the length of the projector tube. The heat gener- `ated by burning the propellent charge 47 causes ignition of the black powder in the thermal element 55 a predetermined interval after the burning of the propellent 47 has ceased suicient to insure that the projectile is a safe distance away from the projector. The pow- `der in thermal element 55 ignites train S6 which in turn ignites the black powder primer 58 for the primary propellent 43. The gases liberated by the primary propellent build up a pressure sufficient to cause the plates 52 to be bent rearwardly of the combustion chamber at the tab portion 53 to uncover the openings and thereby permit exit of the gases into the second chamber for high velocity discharge through said nozzle.

I claim:

1. A rocket projectile comprising a body having a substantially cylindrical bore therein forming a main combustion chamber and an auxiliary combustion chamber, said auxiliary chamber located rearwardly of the said main chamber and having a rearward outwardly flaring nozzle, a flight initiating propellant combustible to generate a gas under pressure supported Within said auxiliary chamber, a sustained flight propellant combustible to generate a gas under pressure supported within said main chamber, means to ignite said tlight initiating propellant, a plurality of valves for sealing said main charnber from the gas pressure generated in the said auxiliary combustion chamber during combustion of the said flight initiating propellant, and means to ignite said sustained ight propellant a predetermined time after the said ignition of the said ight initiating; propellant, the said valves movable upon combustion of the sustained tlight propellant to interconnect the said main chamber with the said auxiliary chamber, the said valves comprising a series of plates disposed in a plane at right angles to the longitudinal axis of the said body.

2. The combination defined in claim 1 characterized by the fact that the said plates are supported in the said plane by an integral and deformable tab element.

References Cited inthe ile of this patent UNITED STATES PATENTS 1,102,653 Goddard July 7, 1914 1,823,378 Scardone Sept. 15, 1931 1,901,852 Stolfa et al. Mar. 14, 1933 2,391,865 Chandler Jan. 1, 1946 2,412,134 Eksergian Dec. 3, 1946 2,412,173 Pope Dec. 3, 1946 2,497,888 Hirschfelder Feb. 21, 1950 2,524,591 Chandler Oct. 3, 1950 FOREIGN PATENTS 379,664 Italy Apr. 2, 1940

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Cited By (34)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2871658A (en) * 1957-12-23 1959-02-03 William J Keck Sustainer exhaust gas deflector
US2879955A (en) * 1951-08-02 1959-03-31 Zborowski Helmut P G A R Von Airborne bodies and in particular self propelled missiles
US2912820A (en) * 1953-07-31 1959-11-17 Quentin R Whitmore Combined ram jet and rocket engine
US2923126A (en) * 1955-02-17 1960-02-02 Soc Tech De Rech Ind Propulsion system
US2954947A (en) * 1958-11-21 1960-10-04 Richard J Zabelka Rocket assisted pilot ejection catapult
US2956401A (en) * 1959-06-12 1960-10-18 Ernest M Kane Variable thrust rocket motor
US2988877A (en) * 1957-12-30 1961-06-20 Phillips Petroleum Co Solid propellant rocket motor
US3000306A (en) * 1958-01-09 1961-09-19 Gen Dynamics Corp Solid propellant propulsion system
US3008414A (en) * 1954-01-21 1961-11-14 Hotchkiss Brandt Self-propelled projectile
US3013494A (en) * 1957-08-09 1961-12-19 Chanut Pierre Louis Jean Guided missile
US3035796A (en) * 1958-11-21 1962-05-22 Cecil A Glass Dual thrust rocket booster tube
US3085511A (en) * 1959-04-22 1963-04-16 Hans O Donner Tail of mortar projectile
US3086359A (en) * 1960-07-19 1963-04-23 Davis Edward James Integral nozzle separator for a multistage reaction motor
DE1147804B (en) * 1959-05-18 1963-04-25 United Aircraft Corp Solid rocket with two series-connected combustion chambers
US3088273A (en) * 1960-01-18 1963-05-07 United Aircraft Corp Solid propellant rocket
US3093964A (en) * 1960-12-14 1963-06-18 United Aircraft Corp Two-stage rocket
US3107487A (en) * 1960-08-12 1963-10-22 Aerojet General Co Rocket motor
US3128600A (en) * 1960-05-18 1964-04-14 Thiokol Chemical Corp Multilevel solid propellant rocket motor
DE1180578B (en) * 1960-09-15 1964-10-29 Ici Ltd rocket propulsion
US3158061A (en) * 1962-10-29 1964-11-24 Samuel E Lager Low toxicity rocket motor
US3169003A (en) * 1963-06-18 1965-02-09 Cecil A Glass Ejection seat apparatus
US3205820A (en) * 1960-03-08 1965-09-14 Jr William C Mccorkle Drag-compensated missile
US3236317A (en) * 1962-07-02 1966-02-22 Dresser Ind Projectile propelling apparatus for use in high temperature environment
US3269313A (en) * 1965-01-18 1966-08-30 William G Willmann Self-propelled sub-munition
US3305194A (en) * 1960-03-08 1967-02-21 Robert G Conard Wind-insensitive missile
FR2167266A1 (en) * 1972-01-11 1973-08-24 Godorazh Georgy
FR2177880A1 (en) * 1972-03-25 1973-11-09 Dynamit Nobel Ag
FR2336560A1 (en) * 1975-12-23 1977-07-22 Imp Metal Ind Kynoch Ltd Destructible retaining member, intended in particular for separation of two pressure chambers of a rocket engine
US4250705A (en) * 1976-12-28 1981-02-17 Societe Nationale Des Poudres Et Explosifs Apparatus for the connection between two stages of a self-propelled engine
US4819426A (en) * 1987-05-08 1989-04-11 Morton Thiokol, Inc. Rocket propelled vehicle forward end control method and apparatus
GB2402462B (en) * 2000-10-31 2005-08-31 Saab Ab Method and device for a multiple step rocket
RU2473819C1 (en) * 2011-11-08 2013-01-27 Открытое акционерное общество "Машиностроительное конструкторское бюро "Искра" имени Ивана Ивановича Картукова" Engine system of safety shutdown system
EP3578791A1 (en) * 2018-06-05 2019-12-11 Diehl Defence GmbH & Co. KG Method for two-step combustion of a solid fuel rocket engine and solid fuel rocket engine
RU2715447C2 (en) * 2017-12-08 2020-02-28 Акционерное общество "Машиностроительное конструкторское бюро "Искра" имени Ивана Ивановича Картукова" (АО "МКБ "Искра") Gas flow flare reflector of solid-propellant rocket engine

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US1102653A (en) * 1913-10-01 1914-07-07 Robert H Goddard Rocket apparatus.
US1823378A (en) * 1929-04-05 1931-09-15 Scardone Charles Rocket
US1901852A (en) * 1930-07-28 1933-03-14 Stolfa Hermann Rocket
US2391865A (en) * 1942-02-14 1946-01-01 Edward F Chandler Self-propelled projectile
US2412134A (en) * 1944-02-12 1946-12-03 Nasa Projectile
US2412173A (en) * 1944-02-22 1946-12-03 Winslow B Pope Projectile
US2497888A (en) * 1944-10-11 1950-02-21 Joseph O Hirschfelder Means for preventing excessive combustion pressure in rocket motors
US2524591A (en) * 1944-07-19 1950-10-03 Edward F Chandler Rocket projectile

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Publication number Priority date Publication date Assignee Title
US1102653A (en) * 1913-10-01 1914-07-07 Robert H Goddard Rocket apparatus.
US1823378A (en) * 1929-04-05 1931-09-15 Scardone Charles Rocket
US1901852A (en) * 1930-07-28 1933-03-14 Stolfa Hermann Rocket
US2391865A (en) * 1942-02-14 1946-01-01 Edward F Chandler Self-propelled projectile
US2412134A (en) * 1944-02-12 1946-12-03 Nasa Projectile
US2412173A (en) * 1944-02-22 1946-12-03 Winslow B Pope Projectile
US2524591A (en) * 1944-07-19 1950-10-03 Edward F Chandler Rocket projectile
US2497888A (en) * 1944-10-11 1950-02-21 Joseph O Hirschfelder Means for preventing excessive combustion pressure in rocket motors

Cited By (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2879955A (en) * 1951-08-02 1959-03-31 Zborowski Helmut P G A R Von Airborne bodies and in particular self propelled missiles
US2912820A (en) * 1953-07-31 1959-11-17 Quentin R Whitmore Combined ram jet and rocket engine
US3008414A (en) * 1954-01-21 1961-11-14 Hotchkiss Brandt Self-propelled projectile
US2923126A (en) * 1955-02-17 1960-02-02 Soc Tech De Rech Ind Propulsion system
US3013494A (en) * 1957-08-09 1961-12-19 Chanut Pierre Louis Jean Guided missile
US2871658A (en) * 1957-12-23 1959-02-03 William J Keck Sustainer exhaust gas deflector
US2988877A (en) * 1957-12-30 1961-06-20 Phillips Petroleum Co Solid propellant rocket motor
US3000306A (en) * 1958-01-09 1961-09-19 Gen Dynamics Corp Solid propellant propulsion system
US2954947A (en) * 1958-11-21 1960-10-04 Richard J Zabelka Rocket assisted pilot ejection catapult
US3035796A (en) * 1958-11-21 1962-05-22 Cecil A Glass Dual thrust rocket booster tube
US3085511A (en) * 1959-04-22 1963-04-16 Hans O Donner Tail of mortar projectile
DE1147804B (en) * 1959-05-18 1963-04-25 United Aircraft Corp Solid rocket with two series-connected combustion chambers
US2956401A (en) * 1959-06-12 1960-10-18 Ernest M Kane Variable thrust rocket motor
US3088273A (en) * 1960-01-18 1963-05-07 United Aircraft Corp Solid propellant rocket
US3305194A (en) * 1960-03-08 1967-02-21 Robert G Conard Wind-insensitive missile
US3205820A (en) * 1960-03-08 1965-09-14 Jr William C Mccorkle Drag-compensated missile
US3128600A (en) * 1960-05-18 1964-04-14 Thiokol Chemical Corp Multilevel solid propellant rocket motor
US3086359A (en) * 1960-07-19 1963-04-23 Davis Edward James Integral nozzle separator for a multistage reaction motor
US3107487A (en) * 1960-08-12 1963-10-22 Aerojet General Co Rocket motor
DE1180578B (en) * 1960-09-15 1964-10-29 Ici Ltd rocket propulsion
US3192708A (en) * 1960-09-15 1965-07-06 Ici Ltd Rocket motors
US3093964A (en) * 1960-12-14 1963-06-18 United Aircraft Corp Two-stage rocket
US3236317A (en) * 1962-07-02 1966-02-22 Dresser Ind Projectile propelling apparatus for use in high temperature environment
US3158061A (en) * 1962-10-29 1964-11-24 Samuel E Lager Low toxicity rocket motor
US3169003A (en) * 1963-06-18 1965-02-09 Cecil A Glass Ejection seat apparatus
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