US3044255A - Powder propulsive for rockets or other self-propelled projectiles - Google Patents

Powder propulsive for rockets or other self-propelled projectiles Download PDF

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US3044255A
US3044255A US506766A US50676655A US3044255A US 3044255 A US3044255 A US 3044255A US 506766 A US506766 A US 506766A US 50676655 A US50676655 A US 50676655A US 3044255 A US3044255 A US 3044255A
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charge
rocket
chamber
combustion
diameter
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US506766A
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Precoul Michel
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SOC TECH DE RECH IND
Technique De Recherches Industrielles Et Mecaniques Strim Ste
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SOC TECH DE RECH IND
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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
    • F02K9/08Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/10Shape or structure of solid propellant charges
    • F02K9/18Shape or structure of solid propellant charges of the internal-burning type having a star or like shaped internal cavity
    • 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

Description

rates 3,044,255 Patented July 17, 1962 3 044 255 POWDER PROPULSV FR RUCKETS 0R STEER SELF-PROPELLED PROJECTILES Michel Preconl, Paris, France, assignor to Societe Technique de Recherches industrielles et Mecaniques S.T.R.I.M., Paris, France, a society of France Filed May 9, 1955, Ser. No. 506,766 Claims priority, appiication France May 14, 1954 4 Claims. (Cl. Sti-35.6)
This invention relates to rockets and similar missiles wherein the combustion of an explosive or combustible charge carried by the missile serves to generate a jet of exhaust gases which is discharged through an exhaust nozzle to provide a forward thrust operative toV propel the missile along its path.
It is a general object of the invention to provide an improved propulsive arrangement for a device of the type specified.. Another general object'is to provide an improved rocket construction.
A more specific object is the provision of an arrangement lwhereby the available space alloted to the vcombustible charge in a rocket propulsive system can be more completely illed to capacity than heretofore w-ithout interfering with a satisfactory combustion of Said charge; it is an object consequently toprovide a rocket capable of being provided with a more powerful propulsive charge than was heretofore possible for a rocket of similar size. Another object is to improve the uniformity of combustion of the rocket charge and thereby impart la more uniform thrust, with consequent increase in the accuracy of fire.
A further object is the provision of `an improved firing or detonating arrangements for a rocket; a specific object is the provision of a single firing device for both the front and rear combustible charges of a two-charge rocket. Other objects include the provision of improved centering means for the rocket charges, and improvement of the flow of combustion gases through the rocket. An important object resides in the provisionof a relative arrangement and design of a front combustible charge and a rear 'combustible charge in a rocket assembly, whereby controlled combustion of both charges if automatically obtained, improving the flight characteristics of the rocket on its path.
According to van important aspect of the invention, there is provided a rocket including a front combustible charge having an axialduct therein, a rear combustible charge axially spaced from the front charge and means defining a continuous annular space around said rear charge for the free ow of combustion gases therearound l towards Va rocket exhaust, and means for tiring said charges, whereby owing to the provision of said axial duct in the front charge and said annular peripheral space around said rear charge, combustion will concurrently proceed radially outwards from said axial duct in the front charge and radially inwards from said annular space in the rear charge. "I'he above and further objects, features and advantages of the invention will appear as the description proceeds. In the accompanying diagrammatic drawings which illustrate an exemplary embodiment of the invention, v
FIG. 1 is an axial cross section, withV parts broken away, of a rocket embodying improvements according to the invention, and
FIG.` 2 is a fragmentary sectional view, on an enlarged scale, of a preferred embodiment of one detail of the rocket -illustrated in FIG. l. Y
The exemplary rocket illustrated essentially comprises a headsection T having a terminal fuse T1; a propellent section P and an expandable tail structure-E.
'Ihe propellent section P comprises two chambers. A front chamber C is defined by a front cylindrical casing section V and a rear chamber C' is defined by a rear cylindrical casing section V. The two casing sections V and V' are assembled together in any suitable way, such A as by the threaded connection shown. Between the chambers C and C and in the region of the assembly between the two casing sections, there is defined an intermediate chamber Co. As shown, the rear casing section V is somewhat restricted in diameter in the part where it is assembled with the front casing section V; moreover the front chamber is slightly smaller in diameter than the rear chamber. Designatingby D, D and D0 the diameters -of the front chamber, the rear chamber and the intermediate chamber respectively, the following relation therefore holds: D D D.
An improved explosive charge is provided according to the invention, comprising two blocks of explosive arranged for so-called compensated combustion. The front charge consists of a tubular block B formed with an axial duct 1 therethrough and coated with a shell of suitable combustion inhibiting material around its periphery Z and its end faces 3 and 4. The rear charge B1 is a solid cylinder having a diameter D" less than the rear chamber diameter D and formed with suitably proliled convex end surfaces, such as the hemispherical surfaces 5 and 6 shown.
The outer diameter D" of the front charge B is substantially equal to the inner diameter D of the front ,chamber C. Under these conditions it will be observed that the combustion of the front charge B can be effected only through the axial duct 1 and there ywill be no iiow of discharged combustion gases along the outer periphery of the charge. Thus the inhibitor shell surrounding the charge will not tend to be ripped olf or partially separated from around the charge and safe operation is promoted. The front charge B is separated from the warhead T by a transverse partition or bulwark 7 against which the forward end of the charge P is directly applied through the forward end 4 of the combustion inhibitor shell. Owing to the resulting absence of any chamber in the front part of the charge in which pressure might build up, the buildup of dangerously high pressures is prevented and safety is further enhanced.
Interposed Within the casing between the front and rear charges B land B' and extending substantially within the intermediate chamber C0 is a spacer structure comprising a spider 8 which may have substantially the shape of a cross in transverse section. Each arm of the spider is formed adjacent its forward end with radial enlargements or shoulders 8 positioned by rearward abutment against the shoulders defined by the front ends of the rear casing section V. interposed between the perpeni dicular forward ends of the spider arms and the adjacent rear face of the front charge Bare a set of resilient shims vor spring washers 9. The rearwardly extending arms of spider 8 are arcuately contoured to conform with the partspherical front end of the rear charge B so as to support and position the latter. l
The front parts of the spider arms are formed with cut-outs radially inwards of the outer enlargements -8 thereof so'as todeiine a recess positioned within the intermediate chamber CO and adapted to receive therein a novel primer or detonator assembly 10 which in accord'- `ance with the invention constitutes a common detonator for the entire 'explosive charge of the rocket'. The detonator structure will be more fully described presently.
The rear lend yof rear charge B is Vsupported .and centeredfwithin a rear, spider structurell also crossshaped in transverse section and having suitably arcuatecontoured front 4arm surfaces. A centering pin 16 projects axially in a forward direction from the center of the rear spider into a complementary axial recess in the rear end of rear charge B'. Similarly a centering pin 17 projects axially rearwards from the center of front spider S into a recess in the front end of rear charge C. Both end recesses of the rear charge are coated with combustion-inhibiting material.
The exhaust nozzle assembly for the rocket comprises a converging section 12 followed by a diverging outlet section 13 secured to section 12 by a threaded connection. The converging section 12 is provided by the rear casing V of Ithe rocket body and comprises a circular forward section -`12' followed `by a conical section 12 followed in turn by an inflected connecting or throat section 12". The arms of rear spider 11 are rearwardly contoured to conform with the inner contour of the converging nozzle section 12 so as to support and position said spider therein.
Surrounding the exhaust nozzle structure is a cylindrical tail support which constitutes a rear-ward extension of the main body of the rocket and supporting the retractible tail fins E pivoted thereto in any suitable manner so shown.
The detonator charge mentioned above has a forwardly projecting part -10 thereof extending into the axial duct in the front main charge B. imbedded in the midst of the detonator charge 10 is an electric sparking or detonator head from which a cable conductor 14 extends out of the detonator 10 and through the annular space between the outer surface of rear charge B and the inner surface of rear casing section V. The cable 14 is brought out of the rocket through the exhaust nozzle thereof and is supported in the throat section of the nozzle by a fran,gible sealing disc or element 15. The sealing element 15 has la main transverse section and a cylindrical flange section which is clamped between the mating surfaces of the converging nozzle section 12 and diverging nozzle -section 13. Preferably the frangible element 15 has an axial boss 15 formed with an aperture through which the cable 14 is threaded. The boss 15 increases the rigidi-ty of the wall r15 and results in a cleaner breakage of the wall i5 on detonation of the charge, in that it causes the breakage to occur substantially solely by shear action and without bending or radial tearing effects which otherwise would lead to an incomplete destruction of the element 15 and would allow residual parts thereof to remain in the nozzle throat and subsequently impede the outrush of the exhaust gases.
A preferred construction of the shear Wall 15 is partly illustrated in the enlarged cross section of FIG. 2. In this construction the element 15 has a thin shear-able section of thickness e in kits radially outward part and a thick main section 15a provided by a rearwardly directed -boss of thickness e. In FIG. 2, 15jc designates the diameter on which the wall 15 is clamped between the adjacent parts 15b and 15C of the nozzle sections, and 15e designates the outer diameter of Ithe enlargement or 'boss 15a. It will be noted that 15e is -only very slightly smaller than 15f, lso that there is defined a thin shear section which is comparatively very small both in radial extent and in axial or thickness dimension, adjacent to the clamped outer section of the element 15. In one practical embodiment of the invention, the difference between diameters 15e and 15]L was about ll mm., and the enlarged thickness e was about twice the shear thickness e. Such construction results in a particularly clean breakage of the frangible seal `15 and permits accurate calibration of the breaking pressure under which the seal will burst. Moreover, it will be noted that on breakage of the seal 15 all `around the circumference of the shear section (between diameters 15e and 151), the cylindrical surface of shoulder 15g outwardly defining the boss 15a will be immediately guided by 4the closely adjacent wall of the rear nozzle section and quickly and completely expelled out of the nozzle.
It will be noted that the general construction of the rocket described, and the arrangement of the main explosive charges thereinis such as to eliminate the presi f ence of a forward combustion chamber within the front charge B. This in turn eliminates the necessity of providing a supporting spider for supporting the front end of the front charge B, as was generally required in prior rockets of comparable type. 4
In operation, energizing the electric firing device through cable 14 fires the single detonator charge 110. The projecting appendage 10 immediately sets lire to the front charge B in the inner peripheral area thereof surrounding the axial duct in the charge. The front charge B then proceeds to burn radially outwards around its axial duct in a perfectly smooth concentric manner so that it remains centered in the casing throughout its combustion by its outer periphery and end faces surrounded by the inhibitor shell.
Simultaneously or practically so, the flame from the detonator charge is swept rearwards and sets fire to the rear charge B. In this instance the combustion is initiated over the peripheral surface of the charge owing to lthe annular gap present between said surface and the wall of casing V, and proceeds radially inwards therefrom. Throughout the Icombustion the rear charge remains centered by the action of the spiders 8 and 11 and axial centering pins 16 and 17.
'It has been found that the :arrangement of front and rear charges disclosed herein achieves a remarkably uniform thrust throughout the combustion period of both charges, and especially during the critical initial phase of the combustion. At the start of the combustion period the volume flow of gases emanating from the front charge is small owing to the reduced combustion area corresponding to the recessed axial region of the front charge, having the relatively small diameter d. The annular space 18 'around the rear charge is so dimensioned as to afford an appropriate flow section for these gases. Thus this annular space 18 can be provided relatively small, thereby increasing the permissible power capacity of the missile for a given volume. As the combustion progresses, working radially outwards from the axial duct in the front charge and radially inwards from the periphery of the rear charge, the volume flow of gases from the front charge gradually increases and the gas flow from the rear charge concurrently decreases at a comparable rate. Since the flow area defined between the periphery of the rear charge and the casing wall gradually increases as the combustion proceeds, the increasing rate of discharge of gases from the front charge can be adequately taken care of in spite of the small initial dimension of the gap 18, without creating dangerous or objectionable pressure surges liable to detract from the safety and/ or accuracy of fire. Thus it will be seen that owing to the compensated arrangement of lthe charges according to the invention, uniform, safe and reliable operation of the rocket is achieved while -at the same time attaining maximum capacity and filling ratio, i.e. optimum utilization of the available space within the device.
The operation is further enhanced owing to the absence Aof axial fiow of gases through the rear section of the rocket, except possibly during the final phase of the combustion.
As already mentioned, the diameter D of the rear charge is selected only slightly smaller than the diameter D of the front charge.
Assuming the effective lengths of the front and rear charges are equal, it is evident that in order for both charge to burn for the same period of time, lthe condition is D=D"-d, wherev D and D" are the outer dameters of the rear Iand front charges respectively, and d is the diameter of the axial duct in the front charge. I have found it preferable however to give the rear charge a slightly larger value than that satisfying the above equation, so that it will burn somewhat longer than the front chargeand will remain centered by the centering pins 16 and i7 to the last, and prevent flow of the gases along the axis of the rear section of the rocket. While it 5 is true that after the front charge has been burned up, the residual axial plug of combustible material still present in the rear rocket chamber is liable to vbe expelled withou-t burning owing to the sharp drop in pressure which will occur, still itis found that the loss' of the small amount of fuel represented by this resi-dual plug, is more than olset by the advantage the arrangement brings with it.
Otherwise stated, the diameter of the above-mentioned residual plug is substantially D-(.Dd), and its deliberate loss requires the provision of a correspondingly larger-diameter rear charge; I have found however that the fuel material is utilized to much greater .advantage on the periphery of the charge than axially. y
It will Ibe noted that the domed contour imparted t0 the ends of the rear charge according to the invention greatly promotes the smoothness of the peripheral How of gases throughV the device, minimizing friction, turbulence and the like.
The provision of the central chamber C0 in additio to the ladvantages stemming from thesingle central type of detonation or firing provided by the invention, also facilitates the assembly of the rocket using a two-part casing with a threaded connection between the Vcasing parts in the region of such central chamber.
What I claim is:
` 1. A rocket or similar self-propelled projectile comprising a cylindrical casing having a front chamber, an intermediate chamber Iand a rear chamber, a cylindrical block of powderk forming 4the front charge located in said f front chamber having an axially extending channel therein, means for limiting combustion of said front charge to the surface of'said axial channel, a solid cylindrical block of powder located in said rear chamber and being yspaced from the internal surface of said rear chamber and igniting means for said front and rear charges located in said intermediate chamber so that combustion of said front charge takes place through said axial channel exclusively and combustion of said rear charge takes place along its periphery. Y
y 2. A rocket according to claim 1 wherein `the interior diameter of said front chamber is between the interior diameter of lsaid rear chamber and the interior diameter of said intermediate chamber. y
3. A rocket according to claim 1 wherein the diameter of the rear charge is greater than the difference between the diameter Vof the front charge and the diameter of the axial channel of said front charge.
4. A rocket or similar self-propelled projectile comprising a cylindrical casing having a front chamber, an intermediate chamber and a rear chamber, a cylindrical block gof powder forming the front charge located in said'front References Cited in the le of this patent i UNITED STATES PATENTS 487,628 A Lockhart Dec..6, 1892 1,151,258 Fischer Aug. 24, 1915 1,994,490 Skinner Mar. 19, 1935 2,206,809 Denoix July 2, 1935 2,436,364 McDowell Feb. 17, 1948 2,455,015 Mace et al. Nov. 30, 1948 2,561,670 Miller July 24, 1951 2,605,607 Hickman Aug. 5, 1952 2,691,459 Whitmore Oct. 12, 1954 2,793,492 Sage et al May 28, 1957 FOREIGN PATENTS 831,496 France June 7, 1938 1,004,819 France Dec. 5, 1951 1,058,495 France Nov. 4, 1953 France Apr. 7, 1954
US506766A 1954-05-14 1955-05-09 Powder propulsive for rockets or other self-propelled projectiles Expired - Lifetime US3044255A (en)

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Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3270668A (en) * 1964-12-29 1966-09-06 Atlantic Res Corp Well-treating apparatus
US4807435A (en) * 1985-01-26 1989-02-28 Rheinmetall, Gmbh Air-breathing jet engine
US4819426A (en) * 1987-05-08 1989-04-11 Morton Thiokol, Inc. Rocket propelled vehicle forward end control method and apparatus
US4956971A (en) * 1988-08-03 1990-09-18 Morton Thiokol, Inc. Solid propellant canister loaded multiple pulsed or staged rocket motor
US5070691A (en) * 1988-08-03 1991-12-10 Thiokol Corporation Solid propellant canister loaded multiple pulsed or staged rocket
US7958718B1 (en) * 1989-07-07 2011-06-14 Alliant Techsystems Inc. Precision controlled variable thrust solid propellant rocket motor

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1113612B (en) * 1957-07-11 1961-09-07 Soc Tech De Rech Ind Closing body for the nozzle of recoil engines
GB1120391A (en) * 1957-11-22 1968-07-17 Mini Of Technology Improvements in or relating to rocketwise propelled projectiles
DE1156610B (en) * 1958-11-25 1963-10-31 Ici Ltd Rocket engine with two propellant charges
DE1153942B (en) * 1959-04-08 1963-09-05 Dynamit Nobel Ag Solid propellant rocket
US3088273A (en) * 1960-01-18 1963-05-07 United Aircraft Corp Solid propellant rocket
DE1140407B (en) * 1960-04-30 1962-11-29 Heinrich Klein Dr Ing Pressure equalization for missiles
DE1145859B (en) * 1960-06-04 1963-03-21 Boelkow Entwicklungen Kg Recoil engine for operation with solid fuels
DE1210257B (en) * 1960-07-28 1966-02-03 Gotex Ab Solid rocket engine and method of making the same
DE1162637B (en) * 1961-03-15 1964-02-06 Thiokol Chemical Corp Rocket engine
DE1214052B (en) * 1962-07-19 1966-04-07 Onera (Off Nat Aerospatiale) Solid rocket
FR2448636B1 (en) * 1979-02-12 1986-03-28 Luchaire Sa IMPROVED PROPELLER FOR ROCKETS AND OTHER SELF-PROPELLED MACHINES

Citations (14)

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US487628A (en) * 1892-12-06 Spray-nozzle
US1151258A (en) * 1911-01-12 1915-08-24 Schutte & Koerting Co Oil-burner.
US1994490A (en) * 1934-09-11 1935-03-19 Leslie A Skinner Rocket projectile
FR831496A (en) * 1937-01-04 1938-09-05 Sageb Projectile fitted with a propellant rocket
US2206809A (en) * 1937-06-28 1940-07-02 Sageb Sa Projectile
US2436364A (en) * 1946-01-24 1948-02-17 Dominion Merchants Company Ltd Explosive sealing heads for containers
US2455015A (en) * 1946-01-03 1948-11-30 Aerojet Engineering Corp Means for igniting propellant in rocket motors
US2561670A (en) * 1945-07-30 1951-07-24 Aerojet Engineering Corp Ignitor
FR1004819A (en) * 1947-05-19 1952-04-03 Pyro-technical device
US2605607A (en) * 1944-11-16 1952-08-05 Clarence N Hickman Trap for rocket propellent
FR1058495A (en) * 1952-06-18 1954-03-16 Soc Tech De Rech Ind Jet thruster
US2691459A (en) * 1950-12-14 1954-10-12 Lane Wells Co Disintegrable sealing member
FR1075681A (en) * 1953-03-14 1954-10-19 Autocoussin Dura Improvement in spring or ordinary mattresses, cushions and the like
US2793492A (en) * 1944-11-24 1957-05-28 Bruce H Sage Rocket assembly

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GB594513A (en) * 1940-09-04 1947-11-13 Charles Dennistoun Burney Improvements in or relating to projectiles operating with rocket propulsion
FR1011030A (en) * 1948-11-24 1952-06-18 Self-regulating ignition relay for multi-stage powder rockets

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US487628A (en) * 1892-12-06 Spray-nozzle
US1151258A (en) * 1911-01-12 1915-08-24 Schutte & Koerting Co Oil-burner.
US1994490A (en) * 1934-09-11 1935-03-19 Leslie A Skinner Rocket projectile
FR831496A (en) * 1937-01-04 1938-09-05 Sageb Projectile fitted with a propellant rocket
US2206809A (en) * 1937-06-28 1940-07-02 Sageb Sa Projectile
US2605607A (en) * 1944-11-16 1952-08-05 Clarence N Hickman Trap for rocket propellent
US2793492A (en) * 1944-11-24 1957-05-28 Bruce H Sage Rocket assembly
US2561670A (en) * 1945-07-30 1951-07-24 Aerojet Engineering Corp Ignitor
US2455015A (en) * 1946-01-03 1948-11-30 Aerojet Engineering Corp Means for igniting propellant in rocket motors
US2436364A (en) * 1946-01-24 1948-02-17 Dominion Merchants Company Ltd Explosive sealing heads for containers
FR1004819A (en) * 1947-05-19 1952-04-03 Pyro-technical device
US2691459A (en) * 1950-12-14 1954-10-12 Lane Wells Co Disintegrable sealing member
FR1058495A (en) * 1952-06-18 1954-03-16 Soc Tech De Rech Ind Jet thruster
FR1075681A (en) * 1953-03-14 1954-10-19 Autocoussin Dura Improvement in spring or ordinary mattresses, cushions and the like

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3270668A (en) * 1964-12-29 1966-09-06 Atlantic Res Corp Well-treating apparatus
US4807435A (en) * 1985-01-26 1989-02-28 Rheinmetall, Gmbh Air-breathing jet engine
US4819426A (en) * 1987-05-08 1989-04-11 Morton Thiokol, Inc. Rocket propelled vehicle forward end control method and apparatus
US4956971A (en) * 1988-08-03 1990-09-18 Morton Thiokol, Inc. Solid propellant canister loaded multiple pulsed or staged rocket motor
US5070691A (en) * 1988-08-03 1991-12-10 Thiokol Corporation Solid propellant canister loaded multiple pulsed or staged rocket
US7958718B1 (en) * 1989-07-07 2011-06-14 Alliant Techsystems Inc. Precision controlled variable thrust solid propellant rocket motor

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DE1003516B (en) 1957-02-28

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