US3167016A

US3167016A - Rocket propelled missile

Info

Publication number: US3167016A
Application number: US600859A
Authority: US
Inventors: Czerwinski Waclaw; Chamberlin James Arthur; Woodward Frank Arthur
Original assignee: De Havilland Aircraft of Canada Ltd
Current assignee: De Havilland Aircraft of Canada Ltd
Priority date: 1956-07-30
Filing date: 1956-07-30
Publication date: 1965-01-26
Anticipated expiration: 1982-01-26
Also published as: GB982440A

Description

Jan. 26, 1965 'w. CZERWINSKI ETAL 3,167,016-

ROCKET PROPELLED MISSILE Filed July so, 1956 7 Sheets-Sheet 1 g 0 mm m (D m w (5 m m 00 N I N c- (0 E N N v g. N

INVENTORS W CZERWINSKI JVACHAMBERLIN FAWOODWARD BY g ATTORNEYS w. CZERWlNSKl ETAL 3,167,016

Jan. 26, 1965 ROCKET PROPELLED MISSILE 7 Sheets-Sheet 2 Filed July 50, 1956' NO vm mm mm on mm mm hm @N Wm OmnN N05 mm 5mm Om hm MD. a m w m w m .NIWMF. w

ATTORNEYS Jan. 26, 1965 w. CZERWINSKI ETAL 3,167,016

ROCKET PROPELLED MISSILE Filed July 30, 1956 7 Sheets-Sheet 3 INVENTORS W CZERWINSKI JACHAMBERLIN F .A.WOODWARD ATTORNEYS Jan. 26, 1965 w. CZERWINSKI ETAL 3,167,016

ROCKET PROPELLED MISSILE Filed July 30, 1956 7 Sheets-Sheet 4 INVENTORS WCZERWINSKI J.A.CHAMBERLIN FAWOODWARD BY%; z &

ATTORNEYS Jan. 26, 1965 w. CZERWINSKI ETAL 3,167,015

ROCKET PROPELLED MISSILE Filed July so, 1956 7 Sheets-Sheet 5 INVENTORS W CZERWINSKI JACHAMBERLIN F .AWOODWARD ATTORNEYS 1965 w.'czE'Rw|NsK'| ETAL 3,167,016

ROCKET PROYPQELLED MISSILE 7 Sheets$heet 6 Filed July 30, 1956 Nm vmp mmp 0MP mmw INVENTORS WCZERWINSKI J.A.CHAMBERLIN F.A.WOODWARD ATTORNEYS Jan. 26, 1965 w. CZERWINSKI ETAL ,0

ROCKET PROPELLED MISSILE 7 Sheets-Sheet 7 Filed July 30, 1956 m: Om GNP m mm 09 D: MNF vmF DNF .I1.l|.||

United States Patent 3,167,016 RGCKET IRGPELLED MISSILE Waciaw Czerwinski, James Arthur Chamherlin, and

Frank Arthur Woodward, all of Toronto, Ontario, Canaria, assignors, by mesne assignments, to The De Havil- Band Aircraft of Canada Limited, Downsview, ()ntario,

Canada, a corporation Filed July 30, 1956, Ser. No. 600,859 8 Claims. (Cl. 102-49) This invention relates to missiles and more particularly to an unguided two stage missile propelled by the efilux of gases generated by a rocket motor.

In the past, surface installations either on land or on the sea, have been defended against air launched attacks by the use of various specialized devices, the most common being projectiles propelled through the rifled barrel of a gun directed at the target. Other devices include unguided rocket projectiles, fired either singly or in salvo, and electronically guided missiles which are directed to the vicinity of the target by external aerodynamic control surfaces.

Unguided projectiles have the characteristic of low single shot probability of kill, which has been partially overcome by firing high velocity projectiles fitted with proximity fuses, at a high rate from radar directed guns or projectors. However, in view of the advent of fast modern aircraft which can take extreme evasive action, and of electronically controlled glide bombs and improved aerial torpedoes, it becomes necessary to increase substantially the warhead effectiveness of unguided missiles and to reduce their time of flight if an acceptable kill probability is to be maintained.

Homing guided missile systems have been used with a considerable degree of success. The guided missile overcomes the disadvantage of a long time of flight by continuously correcting its path in relation to the target so that the latters evasive action becomes substantially meaningless. However, such a weapon with its large amount of electronic guidance equipment requires a long period of time to develop and is very expensive to manufacture.

The subject invention embodies an unguided two stage rocket propelled anti-aircraft missile which will achieve an exceptionally high single shot probability of kill. The missile described herein has a warhead weighing in the order of 150 lbs. and travels with a mean velocity of 6,000 ft./sec. over an effective range of up to 6,000 yds. This warhead is about eight times as large as a conventional anti-aircraft shell, and its velocity is about three times as high as that of a conventional anti-aircraft shell. These features make it possible to achieve an acceptable probability of kill against any of the targets previously mentioned, against which conventional defensive armament becomes ineffective. In addition, the rocket of the invention has the special features of spin stabilization, which eliminates the necessity of external fins on the outer casing, and thus simplifies ground handling; furthermore it has a simple mechanical construction combined with a minimum of delicate electronic or mechanical systems, thus reducing the cost as compared to that Of a conventional guided missile.

It is obvious that such a missile will have to be aimed in much the same way as a gun and will require afire control system of similar accuracy. A launching device therefore is provided: it is directed by an automatic electronic fire control system which will not be described herein since it is not part of the invention.

It is an object of the invention to provide a two stage rocket propelled missile having spin stabilization and then, on separation, fin stabilization.

It is another object of the invention to provide a simple 3,167,016 Patented Jan. 26, 1965 but elfective mechanical arrangement whereby the two stages are separated when the propellant charge in the first stage is exhausted.

Another object of the invention is to provide structure associated with both the launcher and the missile which ensures that all parts of the missile leave the guiding surfaces of the launcher at one time in order to minimize misguide due to movement of the launcher during firing.

The foregoing and other objects and advantages of the invention will become apparent from a study of the following specification, taken in conjunction with the accompanying drawings, in which like reference numerals indicate corresponding parts throughout the several views, and in which:

FIGURE 1 is a longitudinal cross-sectional view of a missile constructed in accordance with the invention;

FIGURE 1A is an enlarged longitudinal cross-sectional view of the fore and aft portions of the missile of FIG- URE 1;

FIGURE 2 is a longitudinal, part-elevational and partsectional view of the missile at the initiation of separation of the first and second stage rockets;

FIGURE 2A is an enlarged longitudinal, part-elevational and part-sectional view of the fore and aft portions of the missile at the initiation of the separation of the first and second stage rockets;

FIGURE 3 is a transverse sectional view of the missile taken on the line 33 of FIGURE 1;

FIGURE 4 is a perspective view of the first and second stage rockets of the missile immediately after separation;

FIGURE 5 is a perspective view of a head segment with its associated fairing segment;

FIGURE 6 is a perspective view of a launcher for the missile, one of the barrels of the launcher being broken away to show a missile in firing position in the barrel;.

FIGURE 7 is a longitudinal sectional view through, a barrel of the launcher showing a missile in firing position;

FIGURE 7A is a detail of FIGURE 7 on a larger scale and showing a stop for the missile;

FIGURE 8 is an end elevational view of the launcher of FIGURE 6; and

FIGURE 9 is a view similar to FIGURE 7 but showing the missile emerging from the barrel.

Referring particularly to FIGURES 1, 2, 1A and 2A a missile 20 comprises a first stage 21, and a second stage 22 coaxially disposed within and substantially encased by the first stage. The first stage includes an outer shell or casing 23 of cylindrical cross-section, to the aft portion of which is welded or otherwise secured a housing 24 of similar cylindrical cross-section. The housing is, in the vicinity of its attachment to the shell 23, of slightly larger diameter than the shell 23 and has machined a recess 25 in its inner cylindrical surface adjacent its forward edge. The aft portion of the housing is of greater diameter than the forward portion, to provide a rear supporting ring 26 the main function of which will become apparent later. The ring 26 is provided with tapped holes 27 into which bolts 28 are inserted to attach a nozzle assembly 29 comprising an outer ring 30 and an inner ring 31, the space therebetween constituting an annular convergentdivergent nozzle 32 in which are disposed deflector vanes 33. The vanes 33 deflect the exhaust gases through an angle of approximately 20 from the axis to impart an angular velocity or spin to the rocket. The vanes are so arranged that the thrust is directed toward the centre of gravity of the missile in order to minimize the effects of malalignment due to thrust deviations around the nozzle. The vanes 33 are fast to the

rings

30 and 31 and the vanes and the rings together with a spherical segment 34 which encloses the space within the inner ring 31 constitute a rigid unit.

The propellant charge for the first stage rocket comprises a multiplicity of precast rings or grains which are manufactured from some suitable rapid burning rocket fuel. It will be noted that the inner diameters of the rings of propellant 35 gradually increase from fore to aft where, at the final ring, the grain assembly is held in position by a segmented rear grain support 36 having flanges 37 whereby it is rotatably mounted in the recess 25. A forward grain support 38 serves mainly to retain the propellant charge in position during loading and storage.

The inner wall of the first stage rocket is formed by a slender cylindrical shell 39 which also provides the outer casing of the second stage rocket 22. The second stage rocket is coaxially disposed within the first stage rocket and terminates at its aft and in a convergent-divergent nozzle 40 of circular cross-section which is welded to the shell 39. The discharge portion of the nozzle 40 mates at its outer edge with the inner ring 31 of the nozzle 32 and, together with a split retaining ring 41, forms a gastight joint between the

nozzles

32 and 40.

A fin wrapper 42 is constituted by a body portion 43 to which are attached integral fins 44 arranged in a slight spiral. The fin wrapper is fastened to the nozzle 40 in torque transmitting relationship by means of fore and aft splines 45 and 46'machined externally on the outer surface of the nozzle 40; the splines are complementary to and register with

splines

47 and 48 respectively, which are machined internally at the ends of the body portion 43, Thesplines 46 also engage splines 49 machined on the inner riiig 31 of the first stage nozzle assembly 29 to provide a torque transmitting connection between the two stages. A groove 50 machined in the forward face of the body portion 43 mates with a groove 51 machined in the nozzle 40, the two mating grooves being adapted to receive a corrugated wire spring 52.

Grooves

53 and 54 machinedin the

splines

46 and 48 respectively, receive the split retaining ring 41. Notches 55 provided between the segments of the first stage rear grain support 36 allow the passage of the fins 44 on separation of the two stages.

Near the forward end of the shell 39 (see FIGURES 1A and 2A) is a. break in which a retaining ring 56 is welded in place. Somewhat forwardly of the ring 56 a stepped ring 57 is welded to the exterior of the shell 39 for a purpose that will become apparent hereafter. The extreme forward end of the shell 39 terminates in a forward tube collar 58 which is provided with an integral collet 59 in which is held a forward grain support 60. The grain support 60 positions the second stage propellant charge, which charge is constituted by rings or grains 61. To the outer periphery of a dished annular sealing plate 62 is welded an outer collar 63 having a groove 64 which receives a seal 65. A groove 66 machined in the forward tube collar 58 receives a split retaining ring 67 which holds in position the outer collar 63 and hence the sealing plate 62. An inner collar 69 extends through the hole 68 inthe centre of the sealing plate and it is welded to the sealing plate. From the aft side of the collar 69 projects an igniter 70 for the second stage grain; a warhead 71 is attached to the other side of the collar 69 by means of screw threads.

Referring again to the first stage (see FIGURES 1 and 1A) there is welded to the forward end of the shell 23 a forward ring 72 having, integral with its outer surface, eight winglets 73. Spanning the annular space between the forward ring 72 and the retaining ring 56 are eight head segments generally indicated at 74 (see FIGURES 1A and 5); they are annularly arranged about the retaining ring. Each of the head segments 74 comprises a web 75 having inner and outer

arcuate flange segments

76 and 77 respectively, and radially disposed edges 78 which abut the edges of adjacent webs to form gas tight joints. Sealing

grooves

79 and 80 are machined in the inner and outer flange segments respectively to receive O-

ring seals

81 and 82 respectively. To the forward faces of each of the webs 75 are attached two integral ribs 83 having toe portions 84 which are restrained by the stepped ring 57. A notch 85 on each of the toe portions provides an area of high local stress concentration when suitable loads are applied to the ribs 83. A split retaining ring 86 fits into a groove 87 machined on the forward part of the inner surface of the forward ring 72 and mates witha recess 88 in the forward edges of the annulus constituted by the arcuate flange segments 77; an igniter ring 89 abuts the end of the ring 72. The igniter ring 89, which has high electrical conductivity, is held securely in a mounting ring 90 of U-cross-section and which is made of some heat resistant plastic material having good electrical insulating characteristics. The igniter ring 89 is attached to the end of the ring 72 by means of an integral flanged collar 91 which fits within the forward inner surface of the ring 72 and abuts the split retaining ring 86. Electrical contact between the ring 89 and igniters 92 adjacent the first ring of propellant is made by means of eight leads 93, each one of which is connected to a terminal post 94 on the respective eight head segments 74 and makes contact with the igniters. The

igniters

70 and 92 will not be described in detail since they are of known construction.

Low shock, high heat release, closureless igniters preferably are used for reliability. This type of igniter sprays molten magnesium and aluminum powder intimately into the forward disk of propellant; it is readily available in fully developed form and is known to perform well.

Fairing segments 95, (see FIGURES 2, 2A and 5) when assembled, constitute a fairing which extends from the forward ring 72 to the forward tube collar 58. Each of the fairing segments is provided with a pair of stiffening ribs 96, the inner edges of which abut the outer edges of the opposed ribs 83 of the webs 75; in the ribs 96 are holes 97 through which pass bolts 98 to fasten the fairing segments to the ribs 83.

A warhead assembly 99 (see FIGURE 1A) comprises the warhead 71 having a fragmentable casing which contains a high explosive. The warhead 71 is fastened by screw means to the collar 69 and is contained within a warhead fairing 101 shaped in the form of a truncated cone. Secured to the aft edge of the fairing is a sleeve 1-32; the sleeve includes a face 104 which abuts the ends of the fairing segments 95, and a flange 103 which is encompassed by the split retaining ring 67. The forward end of the fairing 101 is closed by a plate 105 fast to the fairing and having a threaded flange 106. A hole 107 in the plate 105 allows the plug end 108 of a fuse 109 to be screwed into the warhead 71. By the assembly of the collar 69, the warhead 71, the fairing 101 and the fuse 109, this portion of the warhead assembly is secured to the second stage rocket 22. A conical tip fairing 110 has at its base a threaded sleeve 111 for attachment to the flange 106. A miniature radar antenna 112 (see FIGURES 1 and 2) receives signals from the ground control station which are transmitted to the fuse 109 to explode the warhead at the most advantageous position with respect to the target. The fuse 109, the warhead 71 and the

igniters

70 and 92 will not be described in detail since they are well-known to those skilled in the art.

In order to ensure that the missile 20 is properly launched, there is provided a launcher unit 113 (see FIG- URES- 6, 7, 7A, 8 and 9) comprising three separate launching barrels 114 capable of firing three missiles simultaneously or in a prescribed sequence. Each barrel consists of a hollow steel cylinder 115 to the exterior of which are integral stiffening rings 116 which provide circular stability, and also axial flutes 117. In the interior of each barrel (see FIGURES 7 and 9) eight integral missile guide rails 118 are positioned equiangularly around the inner surface. In order to achieve dimensional accuracy, it is proposed that the barrels should be manufactured as centrifugal steel castings. This type of construction not only results in a product having acceptable dimensional tolerances but also permits the inclusion of all the exterior reinforcements and the interior guide rails 118.

The three barrels 114 are joined together to form the launcher unit 113 by welding the abutting flutes 117 of adjoining barrels; this eliminates distortion during welding. The unit 113 is supported on trunnions 119 of a torque box structure 120 consisting of steel plate of appropriate thickness welded to two suitably positioned stiffening rings. Geared elevating arcs 121 are welded to the torque box 120 below each of the trunnions 119. End caps 122 are hingedly mounted on brackets 123; each end cap may be swung from open to closed position by a rod and crank mechanism 124 welded to a common hinge pin 125 and actuated by a hydraulic jack 126. The end caps may be closed to protect empty or loaded barrels against the weather or, when a rocket has misfired, the caps may be closed and the missile ejected by compressed air introduced into the barrel between the missile and the caps. A cutout in the wall of each cylinder 115 (see FIGURES 7 and 7A) receives a missile stop 128 hingedly mounted on a pin 129 and urged inwardly by a leaf spring 130. A spring loaded contact pin 131 positioned immediately forward of one of the missile guide rails 118 is connected to the automatic electronic fire control and delivers an electrical impulse to the igniter ring 89 when it is desired to launch a missile. The trunnions 119 are mounted on a carriage (not shown) which contains suitable azimuth and elevation gearing to position the launcher unit 113 for firing.

Assembly Because of the specific configuration of the two stages of the missile, one arranged within the other, and because of the novel means necessary to effect complete separation, at special sequence of assembly operations is required. The first step includes the filling of the second stage shell 39 with its propellant 61, held in position by the grain support 60; then the dished sealing plate 62 together with its outer collar 63 and its inner collar 69 and the igniter 70 are partially inserted into the forward end of the second stage. The seal 65 then is fitted into the groove 64 provided in the collar 63 and the whole subassembly is completely inserted leaving the groove 66 unobstructed. The split retaining ring 67 finally is placed in the groove 66 to complete this subassembly.

The second step in the assembly of the missile is begun by attaching the rear grain support 36 to the casing 23 of the first stage 21 and filling the casing with the propellant 35. The inner sealing ring 81 then is positioned around the retaining ring 56 and the eight head segments stage sub-assembly is inserted into the first stage sub-assembly and it is pushed as far aft within the first stage sub-assembly as permissible in order to facilitate the installation of the split retaining ring 8-6 in the groove 87. After the ring 86 has been fitted, the second stage sub-assembly is brought forward until the recess 88 in the web 75 of the head segments 74 abuts the split retaining ring 86.

The third step in the assembly of the missile relates to the nozzle section (see FIGURES 1A and 2A). The colrugated wire spring 52 is located in the machined groove 51, and the

splines

47 and 48 on the body portion 43 of the fin wrapper 42 are pushed over the nozzle splines and 46, insuring that the fins 44 slide into the notches of the rear grain support 36. The fin wrapper 42 is pushed far enough forward to depress the corrugated spring 52 and expose the groove 53 in the aft portion of the nozzle 40 whereupon the split retaining ring 41 is in- 6 stalled and the fin wrapper 42 is freed in order to let the face of the groove 54 abut the ring 41. The first stage nozzle system 29 is then positioned at the end of the housing 24 with the inner ring 31 mating with the nozzle 40; the nozzle 29 then is bolted to the housing 24.

The fourth and final step involves the assembly of the extreme forward portion of the missile. The warhead 71 is screwed onto the collar 69 after which the warhead fairing 101 is fitted into place with the flange 103 of the sleeve 192 being a push fit over the split retaining ring 67; by screwing the fuse 109 into the warhead 71, the fairing 101 is held firmly in position. The tip fairing 110 is assembled by screwing the threaded portion of the sleeve 111 onto the flange 106. Then to complete the assembly, the igniter ring 89 is fitted into position, the leads 93 fastened on the face of the webs 75 and the fairing segments 95 placed in position with the bolts 98 fastening them to the ribs 83.

Operation In operation, the end caps 122 of each launching barrel 114 are opened and remain open during firing. The forward end of the missile 20 is inserted in the barrel and the missile is pushed forward until one of the winglets 73 engages the missile stop 128, at which stage the pin 131 makes contact with the igniter ring 89 (see FIGURES 6, 7 and 7A). The fore part of the missile is supported on the inner surface of the cylinder 115 by the winglets 73 while the aft part is supported by the rear ring 26 of the housing 24 on the eight missile guide rails 118. It is intended that the launcher be aimed by radar control in which case the command for firing will be automatic, delivering an electrical impulse to the pin 131. The current will flow through the igniter ring 89 and the leads 93 to each of the contacts on the webs 75 of the head segments 74 to initiate combustion of the first of the grains of propellant 35. The hot gases thus formed will, in passing down the diverging passage between the propellant 35. and the shell 39, ignitg the accelerating compound between the discs of propellant. An inhibitor is provided on the cylindrical surfaces of the discs so that the burning takes place on the faces to provide a large constant area burning surface. The gases generated travel radially inwards and then turn through 90 to pass axially down the diverging passage. The passage is rearwardly divergent in order to accommodate the increasing volume of gas as it progresses rearwardly. In order to prevent breaking up of the propellant under the high axial accelerations encountered, the discs are bonded to the shell 23 and are also supported by the rear grain support 36.

The gases, as they pass rearwardly enter the nozzle system 29 and are ejected to atmosphere through the annular nozzle 32 to produce a thrust on the missile. At the same time, the gases react on the vanes 33 to impart a spin to the missile, which spin acts as the stabilizing force while the missile is in flight. The rotational forces on the first stage rocket are transferred to the second stage by means of the splined connection between the inner ring 31 of the nozzle 32 and the aft portion of the nozzle 40. Because of the high pressure generated by the gas as it is formed in the first stage, high collapsing stresses will be imposed on the shell 39 of the second stage. Therefore, the shell must be of sufficient thickness to withstand the high load; it should not be overlooked however that the second stage grain will aid in withstanding this radial compressive stress. If it is considered that an increased shell thickness would seriously affect design criteria, the second stage may be pressurized immediately prior to use.

As the missile progresses down the launcher, it is supported, as stated above, by the winglets 73 hearing on the inner surface of the cylinder 115 and by the rear ring 26 hearing on the guide rails 118. However, the relationship between the distance from the winglets to the rear ring and the total length of the launcher barrel relative to the length of the guide rails is such that at the instant 7 the winglets emerge from the launcher, the rear ring ceases to bear on the guide rails (see FIGURE 9). This has the effect of completely freeing the entire missile from the launcher at one instant, thus eliminating ballistic dispersion due to tip ofis, transverse acceleration and static unbalance.

The second stage igniter 70 is activated at the same instant as the first stage is fired. However, by means of a time delay fuse, the second stage propellant is not ignited until a predetermined period of time after the firing of the first stage. This period of time is calculated to coincide with the burn-out time of the first stage propellant. When the pressure in the second stage reaches a predetermined value (called the separation pressure), the axial force acting on the spherical segment 34 which encloses the second stage nozzle 40, pushes the whole outer shell 23 aft with sufficient force to break the toe portions 84 of the ribs 83 at the notches 85 (see FIGURE The relative movement of the outer shell and the inner shell will cause a swinging movement or radial dispersion of the head segments 74 and of the fairing segments 95 attached to them (see FIGURES 2 and 2A). The dispersal speed will be augmented by the centrifugal force caused by the rotational velocity of the missile, as well as by the aerodynamic lift acting on the fairing segments. During separation, the eight fins 44 at the aft end of the second stage will pass freely through the notches 55 in the rear grain support. If any difference exists between the relative velocities of the two stages at the instant of separation the rear grain support in free to rotate in the recess 25 in order to obviate any ill efiects due to the difference in relative velocities. Moreover, during separation, the head segments 74 with their associated fairing segments 95 will break away as shown in FIGURE 4.

At the instant of separation, the first stage becomes, in effect, a second launcher. Because of the high angular velocity imparted to the first stage by reason of the gases discharging through the inclined vanes 33, the first stage has a high order of gyro stability and is maintained in a straight path during separation. The second stage is held concentric with the outer shell 23 by the fins 44 at the rear and also by the high gas pressure between the

shells

23 and 39. The fins 44 of the second stage are slightly helical so that the second stage retains the angular velocity imparted to it by the first stage and consequently retains a certain amount of gyro stability. However, because of the high length to diameter ratio of the second stage an additional amount of fin stabilization is required and the fins are provided mainly for this purpose.

After separation, the second stage carrying the warhead 71 continues on its path toward the target. It is not required, nor is it even desirable, that the warhead be exploded by impact with the target and for this reason some form of fusing is provided which will automatically explode the warhead at its optimum position in relation to the target. In the embodiment of the invention described herein, it is proposed to use what is known as a Command Fuse. In this type of fuse, a radar range finder at the launching base constantly measures the distance between the target and the missile and, at the optimum distance, sends out a signal which is received by the antenna 112 and which triggers the fuse. However, it will be understood that other types of fusing may prove advantageous for certain applications, e.g. an infra-red fuse operated by heat generated by the target, a Doppler radar fuse, an accelerometer-type distance measuring fuse, or an electronic self-contained distance measuring fuse.

It is to be understood that the form of the invention herein shown and described is to be taken as a preferred example of the same and that various changes in the shape, size and arrangement of the parts may be resorted to without departing from the spirit of the invention and the scope of the subjoined claims.

What we claim as our invention is:

1. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially disposed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internal diameter bting greater than the external diameter of the shell of the second stage motor so that an annular passage for products of combustion is provided between'the propellant and the second stage motor, and'an annular nozzle at the aft end of the shells to discharge to atmosphere the products of combustion emanating from the passage.

2. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially dis posed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internal diameter being greater than the external diameter of the shell of the second stage motor so that an annular passage for products of combustion is provided between the propellant and the second stage motor, a solid propellant bonded to the inner face of the cylindrical shell of the second'stage motor, the said propellant being of annular cross-section to provide a central passage for .the products of combustion thereof, igniting means for the propellant of the second stage motor, and nozzles at the aft end of the passages to discharge to atmosphere the products of combustion emanating therefrom.-

3. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially disposed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internaldiameter being greater than the external diameter of the shell of the second stage motor so that an annular passage for products of combustion is provided between the propellant and the second stage motor, an annular nozzle at the aft end of the annular passage, a solid propellant bonded to the inner face of the cylindrical shell of the second stage motor, the said propellant being of annular cross-section to provide a central passage for the products of combustion thereof, igniting means for the propellant of the second stage motor, and a nozzle at the aft end of the passage of the second stage motor to discharge to atmosphere the products of combustion emanating therefrom.

4. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially disposed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internal diameter being greater than the external diameter of the shell of the second stage motor so that an annular passage for products of combustion is provided between the propellant and the second stage motor, and nozzle means at the aft end of the first stage motor shell to discharge to atmosphere the products of combustion emanating from the passage.

5. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially dis posed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internal diameter being greater than the external diameter of the shell of the second stage motor so that an annular passage for products of combustion is provided between the propellant and the second stage motor, nozzle means at the aft end of the first stage motor and communicating with the passage for the ejection of the products of combustion of the propellant of the first stage motor to propel the missile and to impart to the missile a rotary motion and thus to provide it with gyroscopic stability, and fin means at the aft end of the shell of the second stage motor to provide aerodynamic stability thereto after separation from the first stage motor.

6. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially disposed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internal diameter being greater than the external diameter of the shell of the second stage motor so that an annular passage for the products of combustion is provided between the propellant and the second stage motor, an annular nozzle at the aft end of the first stage motor and communicating with the passage for the ejection of the products of combustion of said propellant, a series of vanes angularly disposed within the annular nozzle to impart to the missile a rotary motion to provide it with gyroscopic stability before separation of the motors, a nozzle at the aft end of the shell of the second stage motor, and fin means associated with said last named nozzle to provide aerodynamic stability to the second stage motor after separation thereof from the first stage motor.

7. A rocket propelled missile comprising a first stage rocket motor including a cylindrical shell, a second stage rocket motor including a cylindrical shell coaxially disposed within and substantially encased by the first stage motor, the cylindrical shell of the second stage motor being spaced from the cylindrical shell of the first stage motor to define therewith an annular chamber, a solid propellant in the annular chamber and bonded to the inner face of the cylindrical shell of the first stage motor, igniting means for the propellant, the solid propellant being of annular cross-section and its internal diameter being greater than the external diameter of the shell of the second stage motor so that an annular passage for products of combustion is provided between the propellant and the second stage motor, an annular nozzle at the aft end of the annular passage for discharging the products of combustion of the first stage motor to atmosphere, a series of vanes angularly disposed within the nozzle to impart to the missile a rotory motion and thus to provide it with gyroscopic stability before separation of the motors, a nozzle at the aft end of the second stage motor for discharging its products of combustion to atmosphere, and a group of fins equiangularly arranged on the nozzle of the second stage motor to provide the second stage motor with aerodynamic stability after separation of the motors.

8. A missile as claimed in claim 7, in which the fins are helical in shape with their helix axis coinciding with the longitudinal axis of the second stage motor.

References Cited by the Examiner UNITED STATES PATENTS 562,535 6/96 Hurst l025l X 1,102,653 7/14 Goddard 102-49 X 2,091,635 8/ 37 Hayden 102-69 2,246,429 6/ 41 Brandt 102-69 2,478,774 8/49 Meinel 89-1.7 2,611,317 9/52 Africano 10250 2,686,473 8/54 Vogel 102-49 2,701,984 2/55 Terce 10249 2,818,779 1/58 Koeper 89--1.7

SAMUEL FEINBERG, Primary Examiner. SAMUEL BOYD, Examiner.

Claims

1. A ROCKET PROPELLED MISSILE COMPRSING A FIRST STAGE ROCKET MOTOR INCLUDING A CYLINDRICAL SHELL, A SECOND STAGE ROCKET MOTOR INCLUDING A CYLINDRICAL SHELL COAXIALLY DISPOSED WITHIN AND SUBSTANTIALLY ENCASED BY THE FIRST STAGE MOTOR, THE CYLINDRICAL SHELL OF THE SECOND STAGE MOTOR BEING SPACED FROM THE CYLINDRICAL SHELL OF THE FIRST STAGE MOTOR DEFINE THERWITH AN ANNULAR CHAMBER, A SOLID PROPELLANT IN THE ANNULAR CHAMBER AND BONDED TO THE INNER FACE OF THE CYLINDRICAL SHELL OF THE FIRST STAGE MOTOR, IGNITING MEANS FOR THE PROPELLANT, THE SOLID PROPELLANT BEING OF ANNULAR CROSS-SECTION AND ITS INTERNAL DIAMETER BEING GREATER THAN THE EXTERNAL DIAMTERR OF THE SHELL OF THE SECONE STAGE MOTOR SO THAT AN ANNULAR PASSAGE FOR PRODUCTS OF COMBUSTION IS PROVIDED BETWEEN THE PROPELLANT AND THE SECOND STAGE MOTOR, AND AN ANNULAR NOZZLE AT THE AFT END OF THE SHELLS TO DISCHARGE TO ATMOSPHERE THE PRODUCTS OF COMBUSTION EMANATING FROM THE PASSAGE.