US3698321A

US3698321A - Rocket assisted projectile

Info

Publication number: US3698321A
Application number: US872176A
Authority: US
Inventors: Richard H Wall
Original assignee: Thiokol Corp
Current assignee: ATK Launch Systems LLC
Priority date: 1969-10-29
Filing date: 1969-10-29
Publication date: 1972-10-17
Anticipated expiration: 1989-10-17

Abstract

An improved projectile suitable for use in artillery pieces, through preferably in the higher calibres, is disclosed wherein projectile range upon launching by the gun is considerably extended by means of a solid propellant rocket motor in combination therewith which is characterized by internal means for supporting the solid propellant. Ignition of the propellant produces propulsive gases. The motor is operated after exit from the gun barrel and extends the flight range of the projectile.

Description

United States Patent Wall 1 Oct. 17, 11972 [54] ROCKET ASSISTED PROJECTILE 3,176,615 4/1965 De Matthew ..l02/49.7 X I 2,994 359 8/1961 Westbrook et al.....l02/103 X 21 .RhdH.WllHtll,Al. [7 1 memo a e a 2,816,721 12/1957 Taylor ..60/253 x [73] Assignee: Thiokol Chemical Corporation, 3,122,884 3/1964 Grover et al ..102/103 X Br' t 1, Pa.

0 FOREIGN PATENTS OR APPLICATIONS [22] Filed: Oct. 29, 1969 1,168,803 4/1964 Germany ..102/49.3 [21] Appl. No.: 872,176

Primary Examiner-Robert F. Stahl 52 us. (:1. ..102/49.3, 60/255, 102/102, Brennan 2 A A 7 "102/103 A A 511 1m. (:1 ..F42b 13/02 [57] ABSTRACT [58] Field of Search,....l02/34, 34.1, 49.3,49.7, 103; An improved projectile suitable for use in artillery 60/355 R5 255 pieces, through preferably in the higher calibres, is disclosed wherein projectile range upon launching by [56] References Gifted the gun is considerably extended by means of a solid propellant rocket motor in combination therewith UNITED STATES PATENTS which is characterized by internal means for supporting the solid propellant. Ignition of the propellant 2489953 11/1949 Bumgy produces propulsive gases. The motor is operated 2941469 6/1960 Barn after exit from the gun barrel and extends the flight 3,108,433 10/1963 De Fries et al ..102/103 X range Ofthe projecma 2,773,448 12/1956 Jasse ..60/256 X 1,879,579 9/1932 Stolfa et a1 ..102/34 X 9 Claims, 3 Drawing Figures Fig.l

INVENTOR. Richard H. Wal/ PATENTEU 17 I973 3.698321 sumanrs Fig.2

INVENTOR. Richard H. Wa// ROCKET ASSISTED PROJECTILE BACKGROUND OF THE INVENTION Artisans in the field of rocket propelled missiles have for many years sought to combine the capability of the ordinary cannon, or artillery piece operating by means of an explosive charge for launching projectiles and the like, with the steady state propulsion efficiency derived from the burning of a solid propellant in a rocket motor. These devices, known as rocket assisted projectiles, or gun boosted rockets, and with which this invention is concerned, have not, up to now, proved entirely satisfactory nor performed in accordance with expectations. Many reasons have been advanced for this, not the least of which is the failure to develop a propellant grain which can withstand the accelerations experienced during the difficult launching, or boost phase of the projectile flight, or to develop means to protect the grain. Success has heretofore been elusive most often in the past because of propellant grain cracking, or other, similar deleterious effects induced by the high launching acceleration forces to which the propellant is subjected. It is of course quite obvious that the rocket motor propellant would be subjected to very high linear, tangential and radial acceleration forces during the launch of the missile and due to the spinning thereof imparted by the lands or rifling grooves on the inside surfaces of the barrel of the gun which fires the projectile. Accordingly the invention to be hereinafter described provides, by means of an unique internal casing and propellant support therein, a workable and operable rocket assisted projectile, which can be safely launched by means of ordinary field artillery equipment.

SUMMARY OF THE INVENTION This invention relates to rocket motors; more particularly it relates to rocket motors adapted for attachment to a gun boosted, or launched projectile.

It is well known in the artillery art that the range of a gun used to fire projectiles, or other ammunition rounds, can be considerably increased by the use therein of a rocket motor to impart propulsive forces thereto. Such a rocket motor is usually rigidly affixed to the projectile and gas evolved from an ignited propellant provide augmenting propulsion efficiency. Ignition of the rocket motor can be accomplished after leaving the gun barrel with a suitably designed ignition system contained within the casing, or ignition can be accomplished by means of the high temperature gases from the launching charge in the gun barrel. There are, however other problems in the attainment of an operative rocket motor for projectiles which do not find such easy solutions. A primary concern is the very high acceleration forces experienced by such projectiles when launched from a gun since unwanted and difficult problems arise therefrom. With respect to the projectile and its warhead pay load, these problems are of relatively little concern since adequate solutions have long been available. However, with respect to the rocket motor, a different situation obtains. Motor casings of ample strength have long been available; however, many state-of-the-art propellants are not so durable, hence in providing a rocket motor for augmenting the propulsion of the projectile after launching by the gun, the artisan must insure that a propellant grain is ineluded which is capable of withstanding the launching forces, so that upon ignition, smooth combustion, and therefore propulsion is obtained. From the above and other considerations to be hereinafter referred to, it appears quite obvious that a most critical component in a rocket assisted projectile, from the standpoint of satisfactory operation and performance, is the propellant.

In its preferred embodiment the present invention sets forth a solution to the problem above referred to in the form of composite propellant comprised of an oxidizer and an elastomeric binder such as carboxyl terminated polybutadiene in combination with a novel support means in a casing containing the propellant.

Accordingly, it is an object of this invention to pro vide a rocket motor in combination with a gun launched projectile for augmenting propulsion of the projectile after exit thereof from the launching gun.

It is another object of this invention to provide a device of the character described which utilizes a solid propellant grain, contained and supported in the rocket motor casing, so as to successfully operate even under the extreme acceleration forces to which it is subjected during the critical launch period.

A still further object of this invention is to provide a gun launched, or boosted projectile, adapted to be propelled after launching by a solid propellant rocket motor, said motor comprising a casing within which the solid propellant grain is so supported, that it becomes capable of undergoing strenuous accelerations without harmful effect.

Still another object of this invention is to provide a solid propellant rocket motor of the character described which is adaptable to supply post fire propulsion to a gun boosted projectile, wherein a tube, centrally positioned interiorly of the motor, and within a central perforation in the solid propellant grain, supports the grain under stress conditions during launch which would cause the grain to undergo strenuous deformation without such support.

With the above and other objects and advantages in mind as will become apparent to those skilled in the art to which the invention pertains, the present invention comprises a rocket motor adaptable for use with a projectile fired from a gun. The motor includes a casing, an elastomeric solid propellant grain in the casing and a support means therewithin of a generally hollow support tube. The tube communicates with the nozzle of the rocket and serves both to conduct the hot gases produced by the solid propellant grain when ignited and to support the grain during the gun boosted phase of the projectile-motor flight. Therefore these and other objects will hereinafter become more readily apparent from the following description taken with reference to the drawings, wherein like reference numerals refer to like parts throughout, and in which are described the preferred and additional embodiments of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. I is a view, partially in section, illustrating one embodiment of projectile incorporating a rocket motor of the invention;

FIG. 2 is a view similar to FIG. 1 showing the invention in another embodiment illustrative of a propellant configuration suitable for one type of burning, and;

FIG. 3 is a longitudinal sectional view of a portion of another and preferred embodiment of the invention illustrating propellant deformation, under linear (left side) and radial (right side) acceleration stresses and further illustrative of a propellant configuration suitable for another type of burning.

DETAILED DESCRIPTION OF THE INVENTION To achieve a more detailed understanding of one embodiment of the invention herein presented, reference can be had to FIG. 1. FIG. 1 shows a projectile 10, including by attachment thereto, a rocket motor 1 1 and a warhead 12. Motor 11 comprises a casing 13 of steel, or other suitable pressure vessel material such as reinforced fiberglass, or a laminate thereof to obtain a structure capable of withstanding high internal pressures and external forces. Positioned in casing 13 is a solid propellant charge 14 comprised of, preferably, an inorganic oxidizer and an elastomeric binder such as carboxyl terminated polybutadiene. Propellant 14 is mixed and cast into a sleeve, or cartridge 15 made of a resin such as epoxy, phenolic or polyisoprene with an inorganic filler of glass particles, asbestos or titanium dioxide. A liner or insulator 24 of a carbon-filled rubbery material can be interposed between the surfaces of propellant 14 and cartridge 15 to provide bonding, or inhibiting qualities if desired. The exact inhibiting and bonding materials employed in liner or inhibitor 24 relates to the type of burning desired; i.e., end,radial or a combination thereof, and each is contemplated within the scope of the present invention as will be shown hereinafter. In essence, however, the invention is more properly concerned with a novel support means for propellant 14 whereby state-of-the-art propellants can be utilized successfully in gun launch applications. As will also be more fully explained in what follows, the propellant grain 14 is selected and formulated to be capable of readily deforming within limits controlled by a support means without damage during the high acceleration phases of projectile flight. Propellant 14, as stated, can be cast in place in cartridge 15, and after proper cure, the loaded cartridge can be inserted, i.e., cartridge loaded in casing 13. Cartridge 15 is formed with a central tube 16, which provides support for propellant 14 and forms therein a central perforation 17. Support tube 16, as indicated above, serves as a support member for propellant 14 during the launch, or firing stages of projectile 10, and can be formed with side perforations 18 (FIG. 3), which are advantageous in some radial burning operations, or can be, as is usual, formed with solid walls, for reverse end burning operation. Inhibitor 24 is bonded to the surfaces of central perforation 17 but not to the outer surfaces of tube 16 for the following reasons:

a. Propellant 14, during environmental storage over varying temperature ranges is permitted to expand and contract;

b. Ignition of propellant 14 along the surface of central perforation 17 is prevented when used in end burning applications (cigarette fashion), by insulating against contact thereof with (hot) tube 16, or the hot rocket propellant gases that would fill the annular volume therebetween, said volume being that formed when propellant 14 is radially deformed, i.e., displaced from central tube 16 under the influence of high radial acceleration during axial spinning of projectile 111 and motor 11.

Fropellant 14 burning on the end surfaces only is shown by dashline A in FIG. 1. The arrows in FIG. 1 indicated the direction of flow of hot gas produced by combustion of propellant 14, initiated in-flight by suitable means such as generally designated delay means 19 and igniter 20. A nozzle 21, aft mounted in casing 1 1 is aligned with tube 16 and provides a means for exhausting gases generated by burning propellant 14 to provide post launch propulsion to projectile 10. A plug 22 in nozzle 21 prevents explosion gases produced upon firing of the gun from entering the hot gas chamber defined by tube 16, and thereby causing premature ignition of propellant 14.

It is sometimes desired, however, as indicated previously, to utilize the heat of the launching charge gases for ignition of propellant 14. Accordingly a removable orifice plug 22-a provided in nozzle plug 22 which is easily and rapidly dislodged during launch of projectile 10, after which hot gases flow through the orifice 22-h thereat, through tube 16 and ignite propellant 14. Thus it is possible to effect significant weight savings and economies in projectile 10 by use of such means as above described, since delay mechanism 19 and igniter 211 can be entirely eliminated.

Referring now to FIG. 2, there is shown an end burning configuration of propellant grain 14 in casing 13. However, in this embodiment cartridge or sleeve 15 is omitted and tube 16 is a recess or cavity 23 in casing 11 by adhesive, or other well known means. To insure end burning operation of propellant 14, insulative material or inhibitor 24 is applied on all surfaces, as shown, except the head end surface so that upon initiation of ignitor 211, burning occurs as in FIG. 1 embodiment, and as further indicated by the dash line B in FIG. 2. The arrows again illustrate the direction of hot gas flow, and show the reverse flow end burning aspect of propellant 14.

In FIGS. 1 and 2, at the lower portions thereof, casing 13 is provided with a rotation band 25 which engages with lands in the gun barrel to effect spinning of projectile 111 as it passes through the bore thereof. Thus it is readily apparent that during launch, very great accelerations in the longitudinal direction are imparted which diminish rapidly to some small negative value because of aerodynamic drag when projectile 10 leaves the gun barrel. However, while projectile 10 at this time is undergoing no, or minimal linear acceleration, it is still subject to high radial accelerations, and propellant 14 will, in general, deform as indicated in FIG. 2 by dash line C, due to said high radial acceleration and the forces attendent therewith. Thus the end portions of propellant 14 will be compressed during this period and will tend to move upward toward the insulated head end wall 27, and the surfaces of perforation 17 away from tube 16 to provide therewith a generally annular volume or gap 28, which is wider at the head end and narrower in the aft end regions.

Referring now to FIG. 3 there is shown an additional and preferred embodiment of the invention wherein propellant 14 and tube 16 are arranged for radial ignition and subsequent burning thereof.

Grain 14, as shown in FIG. 3 is configured for radial burning, utilizing central tube 16 as in the reverse endburner configuration of FIGS. 1 and 2. The major differences in grain 14 for radial burning and for endburning are in the location and thicknesses of inhibitors or insulation. A critical feature of the radial burner is the provision of gap or annular space 28 between propellant grain 14 perforation 1'7 inner diameter and central tube 16 outside diameter at the time of ignition. Gap 28 can be closed under some conditions of tem perature cycling and normal storage, but when projectile and motor 11 are spinning in operation, there must be provided sufficient free volume (end-play or air space) to allow gap 28 to form. On the other hand, a free volume (in the end of the motor) should be provided only to the extent necessary to form an ignition gap, since too much free volume thereat permits excessive strain in propellant 14 during spinning, causing cracks to form in grain 14. It has been found in most instances that confining gap 28 to a rage of 0.05 to 0.10 inch is ample for ignition to take place in a satisfactory manner.

It is usually not necessary to provide radially burning grain 14 with other than a rather thin layer of inhibitor or insulation because in this configuration burn time is short, hence motor casing 13 is exposed to high heat for a only short time. The inhibitor or liner must be bonded to propellant 14 but it is not necessary, or in some cases even desirable, for it to be bonded to the case. Thus, propellant 14 can be cast into a separate mold and then installed into the motor 11, or it can be cast-in-place therein.

A layer of inhibitor or elastomeric material 29 can be applied to both ends of propellant 14- so that burning takes place primarily on the inside surfaces of perforation 17 and the length remains constant. Since inhibited propellant 14 is almost entirely an internally burning cylinder, a graph of chamber pressure plotted against burning time will show a highly progressive trace or curve, maximum pressure being reached at, or near burn-out.

Because of the progressivity of the pressure-time curve or trace, and the high ratio of maximum to average pressure that would normally be encountered with radial burning configuration of grain 14, an ablative throat insert can be advantageously used in nozzle 21. This can be a molded phenolic resin filled with asbestos, rather than an erosion resistant graphite as usually used in the end-burner configuration. Wlth an ablative or eroding throat, the throat increases in diameter as surface area increases. This permits operation of the motor 11 at a higher average pressure, thereby achieving higher rocket-projectile performance.

While components utilized with radial burner grain 14, such as motor case 13, central tube 16, nozzle closure 22, igniter initiator 19 etc., are the same as those for end burner configuration of FIG. 1 and FIG. 2, igniter 20 for the former is selected to be larger is size because the throat area of nozzle 21 and propellant 14 surface area are larger; however in principle both operate similarly.

In operation, referring to FIG. 3, dash line E is representative of mode of deformation suffered by propellant 14 under linear, or longitudinal acceleration conditions. It is readily seen that a compressive force causes head end surface 2) to move downwardly away from insulating wall 27-A, and also gap 28 is at least in part occupied by deformed propellant 14. As spin rate increases, propellant M responds by deforming as shown by dash line F in the right hand of FIG. 3. In this position it is longitudinally elongated until surface 29 abuts the underside of insulated wall 27A, such that continued generation of centrifugal force, will cause compression of the now elongated propellant 14. Further, a larger portion of core perforation 17 is now exposed and ignition along the surface thereof more easily occurs. Since an inhibitor of sufficient thickness has been applied to surface 27 burning will continue in core 17 in a radial direction therealong. Burning as noted above will be rapid because of the larger burning surfaces, and burn time foreshortened.

Having described the invention and its several embodiments, it is obvious that other embodiments and uses of this invention will occur to skilled persons. Accordingly what has been described as an invention and for which Letters Patent therefor is desired is set forth in the following claims.

What is claimed is:

1. A rocket motor assisted projectile, said motor and projectile launched by a gun the in-flight propulsion means for which is said rocket motor, said motor consisting of a propellant grain having a central perforation and a casing, the inner surface of said casing defining a propellant chamber, said casing terminating at its rear end in at least one nozzle; a supporting tube for said propellant grain, said tube forming the inner wall for said propellant chamber and being connected to said chamber through at least one opening, said grain consisting of a solid propellant having an ultimate strain adapted to withstand relatively high axial and tangential forces when supported by said tube, said supporting tube extending into the central perforation of said grain and having an outer dimension which in a substantial number of cross sections thereof is less than the dimension of the central perforation in said grain, so that the grain during the launch of said projectile from said gun is forced to take support against said tube without cracking, but so that when said axial force ceases said grain returns to its earlier shape and then elongates due to centrifugal force to expose a space between the grain and the supporting tube whereby at least radial burning is rendered possible.

2. A rocket assisted projectile according to claim 1 wherein the length of the grain is shorter than the length of the rocket chamber, such that a space is formed between the front end of the motor and the front of the grain, so that the grain after launch of said projectile and motor, is elongated by centrifugal forces until until contained by the opposite ends of said rocket chamber, and the space between said grain and said supporting tube becomes larger.

3. A rocket assisted projectile according to claim 1 wherein said grain is at least partially bonded to said inner surface of said casing.

4. A rocket assisted projectile according to claim 1 wherein the space between the supporting tube and the grain is substantially uniform.

5. A rocket assisted projectile according to claim 1 further including a nozzle plug for preventing hot gas from said launching gun charge from entering the interior of said propellant chamber.

8. A rocket assisted projectile according to claim 1 wherein the tube material is selected from the group consisting of aluminum, magnesium and alloys thereof.

9. A rocket assisted projectile according to claim 1 wherein thickness of the supporting tube walls is substantially uniform.

Claims

6. A rocket assisted projectile according to claim 5 wherein said plug has a closed orifice and said orifice is adapted to be opened by said gas from said launching gun charge, whereby said gas enters said chamber and ignites said propellant grain.

7. A rocket assisted projectile according to claim 1 wherein the material of the supporting tube is ablatable by the gases produced by said solid propellant grain.