United States Patent: 3554078

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( 55214 of 55214 )

United States Patent	3,554,078
Horvath	January 12, 1971

SPHERICAL MISSILE AND LAUNCHING MEANS THEREFOR

Abstract

A spherical jet-propelled missile and launching device adapted to securely retain and support the spherical missile oriented for stable rotation about a spin axis at a small angle of elevation relative to the straight line path of flight of the missile for a short period of time following ignition and during which time the launching device utilizes means such as a motor or a turbine operated by the hot exhaust gases of the missile to spin the missile about the spin axis and preferably also utilizes the gases to cause release of the missile by thermal activation or melting of a thermally degradable bond by which the missile is initially retained to the launching device.

Inventors:	Horvath; Joseph S. (Riverside, CA)
Appl. No.:	04/797,903
Filed:	February 10, 1969

Current U.S. Class: 89/1.808 ; 102/377; 42/105; 89/1.819

Current International Class: F41F 3/052 (20060101); F41F 3/00 (20060101); F41C 27/06 (20060101); F41C 27/00 (20060101); F41F 3/048 (20060101); F42B 15/00 (20060101); F41f 003/04 ()

Field of Search: 89/1.808,1,1.3 244/3.11 102/49.4,49.5 42/1

References Cited [Referenced By]

U.S. Patent Documents


533572	February 1895	Unge
2114214	April 1938	Damblanc
2730927	January 1956	Bourquard
2945420	July 1960	Babcock
3009393	November 1961	Lynch
3245350	April 1966	Kelly

Primary Examiner: Engle; Samuel W.

Parent Case Text

This application is a continuation-in-part of application Ser. No. 671,284 filed Sept. 28, 1967, now abandoned.

Claims

I claim:

1. Apparatus for facilitating launching of a spin-stabilized spherical jet-propelled missile having an exhaust nozzle, said apparatus comprising:

first means, including rotary means and means for supporting the rotary means for rotation about a spin axis coaxial with the exhaust nozzle and for disposing the spin axis in a plane including a line of sight, said rotary means comprising turbine means and support means including fusible means to support and connect such missile to said turbine means and to conduct exhaust gas from such missile to said turbine means and to transmit torque from said turbine means to such missile incident to receiving exhaust gas from such missile;

said support means including said fusible means having a passage therethrough for conducting gas exhausted by such missile through at least a portion of said fusible means and to said turbine means from such nozzle, said fusible means thus being disposed to be heated by high-temperature exhaust gas expelled by such missile;

whereby during initial stages of flow of exhaust gas from said missile said fusible means transmits torque from said turbine means to such missile to spin the latter and concurrently during spinning said fusible means heats toward fusion temperature and softens and releases such missile for spin-stabilized flight in the direction of said line of sight.

2. Apparatus according to claim 1, in which said turbine means comprises reaction turbine means disposed for rotation about said axis and having radial and tangential passages connecting with said support means to receive exhaust gas to provide torque-creating thrust for spinning and said rotary means during reception of gas.

3. Apparatus according to claim 1, said apparatus including sighting means for sighting and aiming said first means to bring said spin axis into a plane including a line of sight of said sighting means.

4. Apparatus according to claim 1, in which said rotary means comprises means providing three points arranged to contact the spherical surface of a supported missile forwardly of the exhaust nozzle thereof to provide nonredundant stabilizing support for such missile during spinning of the missile prior to release thereof.

5. Apparatus according to claim 4, in which said means providing said three points is of cup shape having an approximately hemispherical recess with a substantially annular rim having three inwardly extending projections providing said points and having a base having an exhaust-gas-receiving passage about said spin axis and having an axial extension from said base with annular bearing surfaces thereon, said first means including bearings in said means for supporting the rotary means .

6. Apparatus for facilitating launching of a spin-stablized substantially spherical jet-propelled missile adapted to rotate about a spin axis, and having an exhaust nozzle from which exhaust gas issues rearwardly to propel the missile forwardly along a path, said apparatus comprising:

first means, including supporting means having bearing means defining a spin axis;

second means, including rotary means rotatable in said bearing means about said spin axis, said rotary means having missile-engaging means including connecting means adjacent said axis rearwardly of said nozzle and including further means engaging a plurality of spaced-apart points on the generally spherical surface of such missile to aid in stabilizing such missile during rotational acceleration of the missile about said axis, said connecting means comprising a heat-softenable means positioned in the path of exhaust gas from said nozzle to thereby become softened to the point of structural failure to release the engaged missile; and

third means, including means for causing rotation of said rotary means to transmit torque through said missile-engaging means to such missile to rotationally accelerate the missile.

7. Apparatus according to claim 6, in which said second means comprises gas turbine means utilizing exhaust gas from a supported missile to produce torque to rotate the rotary support device.

8. Apparatus according to claim 6, in which said second means comprises torque-transmitting means engaging said rotary support device to impart torque thereto to rotationally accelerate said device, and means for creating torque in said torque-transmitting means.

9. Apparatus according to claim 6, in which said rotary support device comprises a cuplike portion presenting a recess for reception of a rearward portion of such missile, said cuplike portion having three equally spaced inwardly extending projections disposed for frictional engagement with the spherical surface of such missile, said support device comprising as said attaching means a threaded portion arranged to receive a threaded portion of the missile.

10. The combination with a spherical spin-stabilized rocket-propelled missile adapted to be rotationally accelerated about a spin-axis for stabilized flight along a flight path forwardly of a launching station, of:

first means, including means defining a spin axis and rotary missile supporting means rotatable therein about said spin axis, said supporting means including means engaging the spherical surface of said missile at a plurality of points spaced away from said axis and forwardly of the rearward extremity of said missile, and said supporting means including torque-transmitting fusible means releasably restraining and securing said missile to said supporting means at the rearward extremity of the missile adjacent to said spin axis and in the path of hot exhaust gas from the missile;

second means, including means to impart torque to said rotary missile supporting means to rotationally accelerate said supporting means and the attached missile to stabilizing rotational speed; and

said first means being arranged to use heat from the exhaust gas of the restrained missile to soften the fusible means to effect release of said missile from said rotary supporting means subsequent to rotational acceleration of the missile.

11. Apparatus for facilitating launching of a spin-stabilized spherical jet-propelled missile along a forwardly extending path, said apparatus comprising:

first means for rotationally accelerating such missile to stabilizing rotational velocity about a spin axis inclined at a slight angle to the forwardly extending path, said first means comprising a rotary missile support device and means defining a spin axis and including means supporting said rotary support device for rotation about said spin axis, said support device comprising heat-softenable torque-transmitting attaching means closely adjacent said spin axis for attachment of such missile thereto adjacent the rearmost extremity of such missile, said attaching means being in the path of hot exhaust gas from an attached missile, said rotary support device further comprising means providing a plurality of missile-contacting means for engaging the surface of such spherical missile at points relatively distant from said spin axis for stabilizing such missile during rotational acceleration;

second means, including exhaust-gas-operated turbine means for causing rotational acceleration of said rotary support device to rotationally accelerate an attached missile to stabilizing rotational speed about said axis; and

whereby said attaching means is effective to absorb heat from the missile exhaust gas and soften to release an attached missile from said rotary support device subsequent to attainment of stabilizing rotational speed by said rotary device and attached missile.

Description

The invention comprises a novel missile-supporting device and a combination of launching device and spherical rocket-motor-propelled missile, novel features of the missile, novel features of the launching device, and novel means for releasably securing the missile to the launching device.

It has heretofore been proposed to obviate the undesirable features associated with a ballistic trajectory ordinarily followed by rockets and like jet-propelled missiles, by forming the missile in spherical configuration, spinning the missile about an axis upwardly inclined relative to the intended straight line path of flight and aligned with the thrust axis of the propulsion jet, and releasing the missile following ignition or activation of the jet. The propulsion is effected by the reaction of the exhaust jet of, for example, the rocket motor, with a part or parts of the motor or missile. Such is disclosed, for example, in the U.S. Pat. to Kelly, No. 3,245,350.

It also has been heretofore proposed, as in the U.S. Pat. to Rae, No. 3,045,596, to launch a spherical jet-propelled missile from a stand, and to control the flight of the missile by selectively controlling auxiliary or control jets by radio or wire connection to the missile from the launching or similar control station, with concurrent utilization of gyro-stabilizing means and other auxiliary devices.

In the case of the "guided" or remotely controlled spherical missile disclosed in U.S. Pat. No. 3,045,596, and similarly controlled spherical missiles, experience has shown that the cost of such missiles, each of which comprises extremely expensive gyro-stabilizer stabilizer and electronic equipment, is prohibitively high, to say naught of the development cost and the expense associated with the complex electronic launching station apparatus.

In the case of the prior art spin-stabilized spherical rocket-motor propelled missiles exemplified by the disclosure comprised in the noted U.S. Pat. No. 3,245,350, difficulty is experienced in stabilizing the missile during attainment of desired rotational speed and in coordinating the spinning, ignition or firing, and release, of the missile. Release of the missile prior to attainment of adequate rotational speed results in undesirable likelihood of unstable flight; and delay of release after attainment of adequate rotational speed results in loss of propulsive effort and range, and requires undesirable prolonged aiming of the launching means. Further, aiming of the patented device is conjectural when the device is used in conjunction with known launcher-holding and aiming devices such as an ordinary rifle. Further, lack of positive-action means to prevent wobbling of the missile during spinup resulted in loss of accuracy in launching the missile along a desired path.

The present invention, by providing simple and inexpensive means for accurately positioning and supporting the missile for stable rotation about a spin axis and means for aiming and quickly and automatically releasing the missile at the time stabilizing rotational speed is attained, obviates all of the aforementioned undesirable features of the prior art spherical rocket missiles and launching devices.

The present invention obviates all of the noted difficulties and undesirable features associated with prior art spherical missiles, and provides a very lightweight and inexpensive combination of elements, by utilizing a novel rotatable missile-supporting and releasing means which nonredundantly contacts the surface of the missile at spaced-apart points on a path or line encircling the spin axis and which utilizes rotary power means such as a motor or a gas turbine means using the hot exhaust gas from the missile, to cause rapid spinup and which uses the hot gas to cause release of the missile. The several means of the invention are such that the path of flight of the missile is along a predetermined line such as the sighting or aiming line of an apparatus, such as a rifle, to which the novel means are attached. In a preferred exemplary form of the invention, the rotatable supporting means comprises a generally cuplike rotary support member which is supported for rotation about a propulsion and spin axis inclined at a selected elevation angle to the predetermined or aiming line, and which member comprises nonredundant supporting means and which is caused to rotate by being driven by power means such as a motor or the like or by a turbine means driven by rocket exhaust gas issuing from the missile following ignition. The exhaust gas in the latter case is caused to produce driving torque on the supporting means, by being diverted into channels or ducts from which it issues in thrust-producing jets from the periphery of the supporting member. The rotary support member, which also transmits torque to the supported missile, comprises connection means permitting connection to the rear or exhaust portion of the spherical missile, and also, as previously indicated comprises a set of three suitably spaced point-supports engaging spaced points on the missile surface forwardly of the exhaust port and preferably equidistant from the spin axis along a circular imaginary line defining a plane transverse of the spin axis.

Release of the spherical rocket missile from the rotary supporting means is best effected by causing hot missile rocket exhaust gas to weaken by heating or to heat and soften or melt a fusible member which prior to weakening by softening or melting, firmly secures the missile to the rotatable support means. The fusible connecting member, which in the preferred embodiment is of the nature of a brazing alloy serving to secure the rocket to the rotary support means, thus also serves to transmit torque from the latter to the missile during the spinup period. The nature of the fusible material and/or other physical characteristics thereof, are so selected as to result in fusion or gross weakening, and release of the missile substantially concurrently with attainment of at least a suitable or desired rate of stabilizing spin or rotation of the missile. And the arrangement of the rotatable supporting means is such that spinup (rotational acceleration to the desired spin rate) is attained in a very small period of time, for example, a small fraction of one second. Further, in an exemplary and preferred embodiment of the invention the rotatable support means including the cuplike member are preferably so formed that the launching device may be attached to a launching and aiming means such as rifle. Further, in a principal form of the preferred embodiment they are preferably so formed that the hot muzzle blast from a cartridge fired in the rifle effects ignition of the propellant charge. The preferred arrangement is, however, such that no part of the missile or launching device is in the path of a projectile or bullet fired from the rifle, or interferes with any supporting means. As will be made evident, other launcher-supporting means may be utilized, especially in the case of large spherical spin-stabilized rocket-propelled missiles. An example of other supporting means is illustrated herein. Also other means for initiating ignition of the propellant charge may be used, such as, for examples, sparks or open flames, and electric igniters.

The previous brief general description of characteristics of the invention make it evident that it is a principal object of the invention to provide general improvements in launching means for stably supporting and means for releasing spin-stabilized spherical jet-propelled missiles.

Another object of the invention is to provide improved means for supporting a spin-stabilized spherical self-propelled missile.

Another object is to provide improved means for supporting and orienting a spin-stabilized spherical missile during spinup and release of the missile.

Another object is to provide improved means for temporarily restraining and automatically releasing a spin-stabilized jet-propelled spherical missile during spinup.

Another object is to provide a novel thermally activated release means for performing the dual function of supporting a spin-stabilized missile during spinup and automatically releasing the missile, utilizing hot exhaust gas from the missile.

Another object of the invention is to provide improved means for temporarily supporting a spherical jet-propelled spin-stabilized missile during rotational spinup thereof.

Other objects and advantages of the invention will be hereinafter set out or made evident in the appended claims and following description of preferred principal and subsidiary exemplary physical embodiments of the invention as illustrated in the accompanying drawings. In the drawings:

FIG. 1 is a pictorial representation of supportive launching means according to the invention, as applied to a rifle, and indicating the angular and coplanar relationship between the thrust and spin axis of the missile and the flight path of the missile;

FIG. 2 is a view partly in section and to larger scale than FIG. 1, depicting a spherical jet-propelled missile and the arrangement of components of the preferred exemplary embodiment of the invention, as applied to a rifle;

FIG. 3 is a partial sectional view taken as indicated by the directors and broken line 3-3 in FIG. 2, showing gas-driven reaction means for spinning the missile, and igniting means;

FIG. 4 is an exploded view, partly in section, showing details of components of an exemplary spherical missile and of an integrated assembly of components including a rocket nozzle, a nozzle support, an expendable member, and a fusible or heat-degradable bond or connecting means;

FIG. 5 is a sectional view of components of the assembly depicted in FIG. 4, taken as indicated by the directors and line 5-5 of FIG. 4;

FIG. 6 is a view of the assembly comprising a fusible member, in disassembled array and in section, showing the separate parts which form a thermally activated release device for temporarily supporting, rotating, and releasing a spin-stabilized spherical missile;

FIG. 7 is a view in elevation, partly in section, showing a spherical missile and a subsidiary form of means for supporting and rotating to spinup speed the spherical spin-stabilized rocket-propelled missile, adapted for use with large machine-carried spherical missiles and for use on machines such as aircraft and armored land vehicles; and

FIG. 8 is a fragmentary view showing details of relations between the spherical missile and rotary- supporting and-retaining means therefor depicted in FIG. 7, viewed as indicated by indicators 8-8 in FIG. 7.

Referring now to the drawings, in FIG. 1 is shown an exemplary spherical rocket-jet-propelled missile 10 supported for rotation about a spin axis A by supporting means denoted generally by ordinal 20, and which supporting means are mounted upon a suitable aiming device in the form of a rifle R. The aiming device, R, is equipped with sighting means such as conventional rifle sights Sr and Sf which define a line of sight, S. The arrangement of the elements including missile 10, supporting means 20 and aiming device R, is such that spin axis A intersects line of sight S and a prolongation of a line of flight F along which missile 10 is adapted to course in stabilized flight, at an angle .theta.; and the structures are further so arranged that angle .theta. is of selected value in accord with principles enunciated in the aforementioned patent to Kelly.

In FIG. 2, the supporting means 20 is indicated, partly in section, as integrally bonded to the lower front end of the rifle or aiming device R. Means 20 comprises a relatively fixed or stationary housing 20h preferably formed integrally with a web 20w which is united along its upper face to the aiming device R and has a shaped portion 20x encircling the muzzle of device R. Secured in housing 20h are bearing means, such as ball bearing units 20b and 20b' in which are rotatably mounted rotary support means which are generally denoted by number 22.

The rotary support means 22x in the exemplary form depicted in FIGS. 1--6, comprises a generally cup-shaped member 22t of cylindrical external shape and which is integral with a rearwardly extending hub and spindle 22s, the latter two portions being dimensioned to fit in the inner races of bearings 20b' and 20b, respectively. Thus member 22t is supported for easy and relatively free rotation about the a spin axis, A, which is defined by the axis of the coaxial bearings 20b and 20b'. The bearings are secured in the housing 20h in any suitable way, as, for example, by retained means 20r and 20p as shown.

The forward end of member 22t is formed with a recess 22r (FIG. 2), slightly larger and deeper than the rear hemisphere of the spherical missile 10 which it is adapted to support. The forward rim 22m of member 22t is provided with three preferably equally-spaced inwardly extending tits or projections such as 22g, all of which are shaped and dimensioned to engage respective small areas or points on the generally spherical surface of missile 10 at or slightly rearwardly of the central circumference when the latter is fully mounted in the support means and thereby provide nonredundant peripheral stabilizing support for the missile and to aid in frictionally transmitting torque to the missile from the member 22t. At the bottom of recess 22r, the cup-shaped member 22t is provided with a tapped bore which is dimensioned and arranged to receive the threaded rear end of a bored expendable member 12 (FIGS. 4 and 6). The bore of member 12 communicates or opens into a shaped chamber 22c (FIGS. 2 and 3), formed in the body of cuplike member 22t. As indicated in FIG. 3, chamber 22c opens radially into each of a plurality (three as shown) of passages 22d, each of which passages terminates at the exterior surface of member 22t in a respective thrust duct or orifice 22d e. Thus gas ejected into chamber 22c via the bore of member 12 is diverted outwardly through passages 22d and orifices 22e where, by reaction, it produces thrust forces which exert torque on the member 22t and cause rotation or spinning of the rotatable structure described, in the manner of a reaction turbine.

The aforementioned bored member 12 (FIGS. 4, 5, and 6) is part of an integrated assembly K which initially is rigid and serves to secure or attach the missile 10 to rotary support member 22t, but which assembly is devised to thermally degrade and part by softening or fusion of a portion thereof and to thereby release the missile at an appropriate determined time. The assembly K is depicted in integrated or unitary form (in Section) at the left in FIG. 4, and the components of the completed assembly are shown broken apart and displaced each from the others in FIG. 6. Included with member 12 in the assembly K is a rocket-nozzle support 10s whose threaded end is adapted to be turned into a complementary threaded bore formed in a plug 10p which is a part of the missile 10. Nozzle support 10s carries a shaped rocket nozzle 10n, which is press fitted into a complementary recess provided in the nozzle support, as indicated.

The assembly K further includes special means adapted to initially firmly and rigidly unite the nozzle support and member 12, but adapted to be grossly weakened by heating whereby to release the nozzle support from member 12. The noted special means, in presently preferred form, is initially in the form of an annulus or ring of fusible alloy 14, such as silver solder or braxing alloy, which is fused to the confronting annular faces of member 12 and nozzle support 10s, and which as indicated in FIG. 4 serves to releasably unite those components of the assembly into a rigid integral unit. In manufacturing the assembly K, the bored nozzle support 10s and member 12 are initially machined to fit in a jig in coaxial alignment, with a fusible alloy preform therebetween. Thus arranged and held in position, the parts are united by fusion of the allow preform by, for example, induction heating, followed by cooling. Thus joined, the nipplelike member 12 and nozzle support 10s have interconnecting coaxial bores as indicated in FIGS. 3 and 4, and are firmly but releasably bonded together by an annulus of the fusible bonding material 14. Following the step of bonding the components into a rigid integral structure or assembly, a set of one or me more (for example, six) small radial bores b (FIG. 5) is provided by drilling through the structure in the region of the fusible bond, whereby to form small radial passages through which hot exhaust gas may escape and thereby accelerate heating of the fusible connection. Such bores may in some instances be formed through only the fusible bonding material, but may, as indicated in the array of disassembled components shown in FIG. 6, be so drilled as to remove portions of one or both of the nozzle support 10s and member 12 as well as portions of the fusible bonding material. The diameters of the bores b, and the number of those bores, depend upon such factors as the composition and melting temperature of the fusible alloy, the radial thickness of the alloy connecting members 10s and 12, and the type of propellant used in the spherical rocket, and are determined for each mark or type of spherical missile 10. As is evident, the fusible bond is made of sufficient strength to withstand handling and missile-spinup stresses and torque.

Following bonding or integration of the nozzle support 10s and member 12 and formation of the radial bores b, the exterior of the assembly K is turned or machined to provide a selected radial thickness of wall at the bond region, and the axially aligned bores are cleaned. Thereafter the nozzle 10n is pressed into place in the nozzle support 10s, and the assembly K is turned home in the plug 10p in the missile. Thus the missile device is provided with a rearwardly protruding temporary adjunct or auxiliary means in the form of member 12 for easy turning into the bore at the bottom of the recess 22r in rotary support 22t. The dimensions of the interfitting components are such that when the missile device thus produced is turned home in the rotary support 22t, the bores b communicate with the open space in recess 22r and with the through-bore exhaust gas passage of assembly K, and the spherical surface of the missile is brought into firm frictional contact with the spaced tits 22q on the interior periphery of the support.

As is indicated in FIG. 2, missile device 10 comprises structural portions which form a compartmented interior and present a substantially spherical exterior surface, and which missile device is dimensioned to be received in recess 22r in firm contact with the three tits or projections 22q. Thus in a forward compartment 10x of the missile device there is situated fuse or detonating means adapted for initiation by deceleration of portions of the device incident to impact of the missile at high speed with a target object. The nature and construction of the means in compartment 10x may vary widely as will be evident to those skilled in the armaments art; and, since per se they are not the present invention they are not herein further described.

In a second compartment 10y of the missile device 10 is disposed explosive charge means adapted to be detonated or exploded by the detonating means housed in compartment 10x. As is evident to those skilled in the art, various types of charge means may be housed in compartment 10y; for example, high-explosive (H.E.) grains or a substantially solid charge of detonatable material.

The chamber-forming means comprised in the structural shell of the missile device further form a third compartment or chamber 10z which is proportioned and disposed in accord with principles well known or stated in the noted patent to Kelly. At the rear this chamber or compartment is walled by means including the aforementioned bored and internally and externally threaded plug 10p (shown in section in FIG. 4), whose external threaded surface engages a complementary threaded surface in the rear interior of the wall or compartment 10z as shown in FIG. 2, and therewith forms a gastight fit or seal. The internally threaded bore 10r of plug 10p is arranged to receive the complementary threaded forward end of the previously described nozzle support 10s as indicated in FIG. 2, the engaging threads or surfaces being made to provide a gastight juncture. The rocket nozzle 10n is tightly sealed in the forward end of the axial bore formed in nozzle support 10s, by the pressure of the press bit, and is seated against a seat formed by an annular step in the bore, as indicated. The nozzle 10n is formed of material resistant to erosion by high-velocity high-temperature gases generated in chamber 10z, and may be, for example, of ATJ graphite. As indicated in FIGS. 2 and 6, nozzle support 10s is provided with an annular enlargement 10s' the forward face of which engages closely a complementary seat formed at the rear of bore 10r of plug 10P, whereby the support is more securely held in the plug and the juncture assured gas-tightness. As shown, the annular enlargement is faced for use of a wrench in turning the assembly K home in the plug 10p.

Thus produced and turned home in rotatable device 22, the missile device 10 is supported and stabilized by contact with tits 22q and is ready for aiming and launching. As practiced with a rifle as indicated in the drawing (FIG. 1), the sights with zero elevation are brought into coincidence with the straight imaginary line of sight S extending to the selected target, and the rifle is discharged using either blank or ball cartridge. In one embodiment of the invention, wherein chamber 22c and passages 22d and 22e are free and open, a portion of the very hot gas moving along the bore of the rifle R under high-pressure passes laterally out through a small lateral port 20z (FIGS. 2 and 3) into one of the ducts 22e and onward through the connecting passage 22d and chamber 22c and through the bore of assembly K and into igniting impingement with charge 10z in the missile. Thus ignited, charge 10z commences burning and produces an exhaust of extremely hot gas under high pressure.

Exhaust gas issues through nozzle 10n into the aligned bores of members 10s and 12 into chamber 22c where it is diverted outwardly through passages 22d into the thrust ducts 22e, from whence it escapes into the ambient atmosphere. A small portion of the gas passes through the relatively small passages or bores b which extend through the fusible material 14 bonding support 10s and member 12 together, from whence it passes into the space between the missile and the bottom of recess 22r between the missile and the cup of support 22t. Gas thus passing into the space between the missile and the bottom of the recess 22r may readily escape into the ambient atmosphere by way of the spaces between projections 22q at the rim of the cuplike member 22t.

As gas commences to escape through thrust ducts 22e, support member 22t commences to rotate about spin axis A. Concurrently gas escapes through small bores b and rapidly heats the fusible bond linking nozzle support 10s and member 12. Within a very short period of time, for example, one-fourth of a second, support 22t and the attached missile 10 attain a rate of rotation sufficient for stabilizing rotation of the missile about the spin axis during flight, and concurrently the fusible bond is heated to above the softening temperature. During the succeeding brief interval of time thereafter, the fusible bond degrades and parts, and the missile is propelled out of the recess in the support 22t by the thrust of the gases being forced through nozzle 10n. The missile rapidly accelerates and travels forward along the line of flight F parallel to and slightly below the line of sight S, leaving expendable member 12 in the support 22t. As is evident, the member 12 may be readily removed by an extracting tool such as is used in extracting broken studs.

To facilitate positioning and retention of a duct 22e in close proximity to port 20z of the timing device R, whereby to insure ignition in the case wherein blast from a fired rifle shell is used, permanent magnet means such as 26 and 27 (FIG. 2) may be utilized, one being secured to housing 20h and the other being cooperatively secured to rotatable member 22t as indicated. As is evident, three magnetic means such as magnet 27 may be utilized on member 22t whereby any one of the thrust ducts 22e may be brought into close proximity to port 20z and retained there by the cooperative action of magnet means 26 and 27. Thus, for example, only a slight manual rotation of the missile and its support 22 need be effected to insure proper positioning for certain ignition by the rifle blast.

In an alternative embodiment of the invention more particularly adapted for use by riflemen in field operations, the missile supporting and stabilizing member 22t is made to be expendable and to be readily inserted and removed from the bearings 20b and 20b'. In that embodiment of the invention and as depicted in FIGS. 2 and 3, the chamber 22c, passages 22d, and thrust ducts 22e are filled with solid propellant composition whereby to further enhance certain and substantially instantaneous ignition of the missile following firing of a rifle. As is evident, the same fusible connection 14 between the missile and support 22t is utilized. Following in ignition and launch of a spinning missile the expendable rotatable supporting member 22t is merely manually withdrawn from the bearings and discarded.

According to the invention, the angular relation of the spin axis A relative to the line of sight S is carefully determined experimentally, in the case of a particular mark of missile, to produce from the missile exhaust a vertical component of acceleration equal to the acceleration due to gravity but of opposite sign; and thus, as explained in the noted patent to Kelly, the missile 10 moves forward along a straight line of flight. Since the latter is offset but a very small distance from (below) the parallel line of sight S, the missile travels toward a point on the target object displaced only the same small distance from the intersection of the line of sight and the target object.

It is evident from the preceding description that the number and size and spatial arrangements of the gas-conducting bores, passages, and ducts may be varied in dependence upon size, mass, and other characteristics of the rotary parts of the missile-launching apparatus and missile, and upon the characteristics of the propellant or propellant gases and of the fusible bonding means or alloy. Also it is made evident that the composition and physical arrangement of the fusible bonding material constituting the thermally activated release means for the missile may be widely varied but should be such that activation by fusion or other weakening during missile launch occurs prior to significant deterioration of any of the other structural parts of the missile and launching means. Further it is made evident that the described means permit rapid and repetitive launches of rocket missiles of the character described, it being necessary only to quickly detach a used rotatable device 22 and replace it with a preassembled unit comprising a missile device 10 and device 22 in those cases where rotary support devices 22 are inexpensive expendable items; or, in the case wherein device 22 is a more permanent part of the launching means, to merely extract the remnant member 12 from device 22 (as with a driver or other tool) followed by turning-home of a new missile device equipped with the described fusible-bond thermal release means and replacement of the spent rifle cartridge.

Additionally it is evident that other means may be used for initiating burning of the propellant charge 10z, as, for example, by fuse means or the like in those instances in which the missile device is large and heavy and is carried by relatively massive launching apparatus transported by carriage or track-laying vehicle.

Launching and supporting means comprising nonexpendable means for rotationally accelerating a spin-stabilized spherical rocket-propelled missile of the character previously described, and especially adapted for use with large missiles of that character, and for all such missiles wherein for reasons of economy in propellant charge or for other reasons it is undesirable to utilize rocket exhaust gases to impart rotational torque for spinup, are illustrated in FIGS. 7 and 8. As was previously indicated, the means for supporting and spinning-up a missile as depicted in the latter drawings are adapted for use with mobile machines such as airborne and land vehicles from which auxiliary power is readily available and which vehicles may carry automatic or remotely controlled aiming supports upon which the missile launching and supporting means may be mounted and thereby aimed. In those drawings the missile 10 and fusible connection 12 are for convenience illustrated as the same as those previously described. The rotary means for releasably holding and stabilizing the missile during spinup and for transmitting spinup torque are similar to those of the first-described embodiment of the invention; but are modified to receive spinup power from external means, and to exhaust the rocket gases to the ambient atmosphere. Thus the rotary support 22t' is formed with a forwardly facing cuplike portion and a set of torque-transmitting and missile stabilizing tits 22q' as before, but the rocket-exhaust gas passages are simpler, and the rearwardly extending stem 22s is arranged to receive applied torque from powered means. Thus the gas exhaust passages to the rear of member 12 comprise one or more radial passages such as 22d', and an axial bore 22x' which may be left open for direct rearward exhaust or may be closed by a plug P whose threaded rear end is received in a threaded enlarged portion of bore 22x' as indicated. Also, the stem or spindle 22s' of the rotary support is, for example, fitted with power-driven means such as a pulley L which receives power through a belt B driven by a motor M. Other equivalent means may be used, such as other power-driven friction means, to rotate the rotary support. As indicated, bearings and a suitable stand or pedestal N supported with motor M on a base Q serve to mount the rotary missile support for rotation about a spin axis whose direction is determined by, for example, a vehicle or movable device to which the base is or may be secured.

As is evident to those skilled in the mechanical armament art, spinup of the missile may be effected prior to an anticipated release of a missile, and the rotation continued, by energization of motor M as a preparatory exercise. At the desired time, the propellant charge may be ignited by any suitable means, including remotely controlled igniter means contained in or replacing plug P. Since spinup may be relatively gradual and the required torque-transmission to the missile effected principally through tits 22q', the fusible link or device connecting member 12 with the missile may be made relatively weak and/or such as to degrade and release the missile when at a relatively low temperature, whereby to very quickly release the missile within an extremely brief interval following propellant ignition. Thus, especially in the case of the modified form of missile-supporting and-releasing means illustrated in FIGS. 7 and 8, the supporting, torque-transmitting, and stabilizing characteristics of the rotary support are important features of the invention.

In the preceding descriptions no specific propellant has been cited or named, since the propellant may be selected from numerous propellants well known in the armament arts, in dependence upon the size, mass, etc., of the missile device 10.

In the light of the preceding detailed description of a preferred and exemplary physical embodiment of the invention, changes within the true spirit and scope of the invention will occur to to others, and hence I do not wish the invention to be restricted to details of the illustrated apparatus and arrangement and mode except as may be required by the appended claims.

* * * * *

Current U.S. Class:	89/1.808 ; 102/377; 42/105; 89/1.819
Current International Class:	F41F 3/052 (20060101); F41F 3/00 (20060101); F41C 27/06 (20060101); F41C 27/00 (20060101); F41F 3/048 (20060101); F42B 15/00 (20060101); F41f 003/04 ()
Field of Search:	89/1.808,1,1.3 244/3.11 102/49.4,49.5 42/1