US3610096A

US3610096A - Spin and fin stabilized rocket

Info

Publication number: US3610096A
Application number: US792958*A
Authority: US
Inventors: Robert E Bauman; James R Christenson
Original assignee: Emerson Electric Co
Current assignee: Emerson Electric Co
Priority date: 1969-01-22
Filing date: 1969-01-22
Publication date: 1971-10-05
Anticipated expiration: 1988-10-05
Also published as: DE1954540B2; IL33056A0; DE1954540C3; JPS491120B1; FR2028841A1; DE1954540A1

Abstract

A ROCKET WEAPON SYSTEM IN WHICH THE ROCKET IS SPIN STABILIZED ALONG ITS FLIGHT PATH BY ROTATION OF THE ROCKET ABOUT ITS LONGITUDINAL AXIS AND IS AERODYNAMICALLY STABILIZED BY A FORWARD FIN RING AND A REARWARD FLARED SKIRT. THE ROCKET IS SPUN-UP IN A LAUNCH TUBE HAVING PLURAL BEARING SUPPORTS AT ONE END FOR SUPPORTING THE REARWARD END OF THE ROCKET, WHILE THE FORWARD END OF THE ROCKET IS SUPPORTED BY FINS ON THE FIN RING, THE LATTER BEING ROTATIONALLY FREE OF THE ROCKET BODY. AFTER SPIN-UP, A BOOST PROPELLANT IS IGNITED TO PROPEL THE ROCKET FROM THE TUBE WITHOUT TIPOFF AND THE USUAL DEBRIS (SABOT FRAGMENTS) ASSOCIATED WITH THE LAUNCH OF SUCH WEAPONS.

Description

oct. s, 1911 R. E. BAUMAN ETAI- sPIN AND FIN STABILIZED ROCKET Ott. 5, 1971 R, `5,.-BNNJMAN E'I'AL 3,610,096

sPIN AND FIN STABILIZED nocxm 2 Sheets-Sheet 2 Filed Jan. 22 1969 United States Patent O U.S. Cl. 89-1.808 13 Claims ABSTRACT F THE DISCLOSURE A rocket weapon system in which the rocket is spin stabilized along its ilight path by rotation of the rocket about its longitudinal axis and is aerodynamically stabilized by a forward iin ring and a rearward flared skirt. The rocket is spun-up in a launch tube having plural bearing supports at one end for supporting the rearward end of the rocket, while the forward end of the rocket is supported by ns on the n ring, the latter being rotationally free of the rocket body. After spin-up, a boost propellant is ignited to propel the rocket from the tube without tipol and the usual debris (sabot fragments) associated with the launch of such weapons.

BACKGROUND OF THE INVENTION 'Ihis invention relates to a weapons system, and more particularly to a weapons system of the rocket projectile type.

Unguided rocket projectiles of the prior art, for use in air-to-air, air-to-ground, ground-to-ground, or ground-toair modes, have been seriously decient in delivery accuracy. For example, in conventional fin stabilized projectile systems, the Weapon is guided along its ilight path solely by the aerodynamic forces exerted by the moving stream of air over the Weapon and its guidance fins. While this type of aerodynamic stabilization tends to maintain the projectile along a predetermined ilight path to its target, it does not compensate for thrust and/or mass misalignment of the projectile itself. That is, common manufacturing diiculties in mass producing such weapons commonly produce off-center thrust forces caused by either nozzle misalignment or uneven burning of the propellant grain, and mass or center of gravity misalignments of the entire projectile. These delciencies cause the projectile to deviate substantially from its intended flight path.

To overcome the above problems associated with fin stabilization, the prior art has developed spin stabilized projectiles which cancel thrust and/ or center of mass misalignments by rotating the projectile about its longitudinal axis during ilight. This tends to equalize such misalignments about the axis of the projectile, thereby producing a canceling effect and enabling the projectile to follow its intended ilight path. Such spin stabilized projectiles cannot, of course, utilize tins for guidance and stabilization because of the tendency of the fins, when exposed to a moving air stream, to reduce rotation of the weapon. Thus, the advantages gained in a iin stabilized weapon system are lost.

Yet another problem commonly associated with weapons of the rocket projectile type is that of tip-off caused .When the projectile leaves its launcher. It is well known that one of the contributing causes of delivery inaccuracy of a rocket projectile is the initial angular deflections and rates imparted to the projectile as it emerges from the launcher. That is, in the time interval just prior to separation of the projectile from the launcher, the rearward or aft end of the projectile is still supported by the launcher after the center of mass of the projectile has passed the forward end of the launcher, and is no longer supported 3,610,096 Patented Oct. 5, 1971 ICC thereby. Thus, the center of mass starts to fall under the influence of gravitational forces imparting a nose-down angular acceleration on the projectile. If the projectile is spinning, a transverse angular rate will be developed because of gyroscopic torques. In the past, attempts have been made to provide the projectile with shoes or sabots which support the projectile on the launching rail or in the launch tube, such shoes or sabots being so constructed so as to simultaneously disengage from the fore and aft rails or the fore and aft portions of the launch tube t0 allow the projectile to be in a free ilight condition as it emerges from the launcher. While such devices have generally proven acceptable for preventing undesirable tipofi, their use has been limited to fin stabilized projectiles. In addition, such shoes or sabots disengage from the rocket after launch and cause debris, or they remain with the rocket and are aerodynamically ineilicient. In most cases, because of their aerodynamic inefficiency, such shoes or sabots are caused to disengage from the projectile after it leaves its launcher. In high-speed, air launched projectiles, the debris caused by the falling shoe or sabot may cause considerable damage to the aircraft launching the projectile, to other aircraft in the immediate vicinity, and to inhabitants and structures on the ground beneath the aircraft.

SUMMARY OF THE INVENTION Accordingly, among the several objects of the present invention may be noted the provision of a rocket projectile having substantially improved delivery accuracy; the provision of such a projectile having the accuracy characteristics of spin stabilized projectiles at low spin rates, thus minimizing design and manufacturing problems associated with high spin; the provision of a projectile of the class described further incorporating aerodynamic stabilization; the provision of a projectile and launcher in which tip-off errors during the launch phase are eliminated; the provision of a projectile and launcher of the class described in which debris normally caused by sabot devices is eliminated; and the provision of a weapons system which is characterized by simplicity ofV construction, low cost, and ease of operation and use.

In general, a projectile constructed in accordance with the present invention comprises an elongate body having a forward portion and a rearward portion. Propulsion means are carried by the rearward portion for propelling the projectile along a llight path to a target, and means are provided for rotating the projectile about its longitudinal axis for stabilizing the projectile during llight. In addition, rotationally free aerodynamic n means are provided on the body, the iin means serving to support the projectile in its launcher prior to launch. Other objects and features will be in part apparent and in part pointed out hereinafter.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a longitudinal View, partly in section, of the rocket projectile and launcher of this invention;

F'IG. 2 is an end view taken on line 2-2- of FIG. l;

FIG. 3 is a section taken on line 3 3 of FIG. 1;

FIG. 4 is a section taken on line 4 4 of FIG. l; and

FIG. 5 is a perspective of the projectile as it emerges from the launcher.

Corresponding reference characters indicate coresponding parts throughout the several views of the drawings.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to the drawings, and particularly to FIG. l, the present weapon system is shown to comprise a rocket projectile 1 within a launcher 3. The projectile 1 is generally comprised of an elongate body having a forward portion 5 and a rearward portion 7, the forward portion carrying a warhead and fuse (not shown) of conventional construction and the rearward portion carrying a propulsion means 9. The propulsion means 9 comprises a rocket boost propellant grain 11 and an exhaust nozzle `13 terminating in a rearward opening 15. Means 17 for rotating the projectile about its longitudinal axis are positioned about the rearward opening 15. The projectile 1 is supported in the launcher 3 by front and rear support means 19 and 2,1, respectively, which also serve to aerodynamically stabilize the missile during flight, as will appear hereinafter.

More specifically, the forward portion S of the rocket projectile 1 is of aerodynamically streamlined configuration and is adapted to contain the warhead and fuse. The rearward body portion 7 of the projectile comprises a propulsion motor case and insulation 25 containing the annular propellant grain 11. The rear support 21 comprises a flared skirt joining the rearward body portion 7 just aft of its midpoint and having an outwardly angularly Haring portion 27 and an integral longitudinally extending portion 29, the latter comprising the rear support for the projectile. The exhaust nozzle 13 is of the converging-diverging type and is affixed to its forward end to the rearward end of the motor case 25 and terminates at its rearward end at the nozzle opening 15.

Within the confines of the diverging portion of the nozzle 13 are the four spin motors 17. These motors are each comprised of a combustion chamber 31 containing a propellant grain 33. A converging-diverging nozzle 35 is in communication with the combustion chamber 31 via a passage 37 extending generally longitudinally from the nozzle to the combustion chamber. The longitudinal axes of the spin motor nozzles 35 extend transversely to the longitudinal axis of the projectile 1 for producing tangential thrust forces which rotate the projectile 1 about its longitudinal axis in a clockwise direction, as viewed in FIG. 2.

The front projectile support is provided by the fin ring 19 which, as best shown in FIGS. l and 4, comprises a ring or band 319 having four radially extending integral fins 41 spaced at 90 intervals around its outer surface. The afin ring 19 is journaled for relative rotation on the rocket projectile 1 between the forward and rearward body portions thereof by means of a bearing assembly 43 interposed between the ring 39 and an inner race 45 carried by the projectile. The yfins `41 and the flared skirt 21 are aerodynamically configured to stabilize the projectile during fiight. In addition, and as will be more fully set forth hereinafter, these fins and the rearward portion 29 of the ared skirt 2,1 support and guide the projectile during the launch phase of its flight.

Referring now to FIGS. l and 3, the launcher 3 is comprised of a cylindrical thin-walled, open-ended, tube 47 having plural supports 49 spaced about its inner periphery extending from near the rear end of the tube to near the center of the tube. As shown in FIG. 3, each of the supports is comprised of an elongate retainer 51 having a dove-tail connection with the interior of the tube 47. A plurality of ball bearings 53 are retained at equal intervals along the length of the retainer with their upper surfaces extending above the upper surface of the retainer. The supports 49 are spaced at 90 intervals around the inner periphery of the launch tube 47 for supporting the aft end of the rocket projectile at its longitudinally extending skirt portion 29. Thus, the effect of the four supports 49 is to decrease the effective internal diameter of the launch tube throughout the extent of the supports. As will be seen, the bearings 53 provide a low-friction support for both longitudinal and rotational movement of the projectile.

The supports 49 and the relative dimensions of the launch tube 47 and projectile 1 are such as to eliminate tip-off during launch. That is, the distance between the vrearward edge of the Ifins 41 and the forward exit end of the launch tube 47 is equal to the distance between the rearward edge of the projectile skirt 21 and the forward edge of the supports 49. In addition, the radial extent or height of the support retainers V51 and bearings 53 is such as to provide sufficient clearance to prevent projectile contact with the launch tube envelope after the skirt 21 clears the supports 49. This clearance, or the difference between the internal fore and aft diameters of the launch tube, is such as to avoid contact under substantially all launch conditions and is sufficient to allow for all perturbations considered to influence projectile-launcher relative displacements, such as projectile mass assymetries, thrust misalignment, initial body angular rates, airloads on the projectile, and launch platform angular and linear rates and accelerations.

Operation is as follows:

The weapons system of the present invention will be described in connection with an air-to-ground unguided rocket projectile, although it should be understood that the weapon may 'be used in other modes, e.g., in air-toair, ground-to-ground, and ground-to-air modes, with either unguided or guided capabilities. In an air-to-ground configuration, the launch tube 47 may be utilized as :both a protective shipping container for the projectile 1 and as a launcher, or only as a reusable launcher. The launch -tube 47 is either movably or rigidly mounted on an aircraft of either the rotary or fixed wing type. In one elnbodiment for mm. rocket projectiles, the tube 4'7 is approximately 48.5 inches in length with fore and aft internal diameters of 6.12 and 5.6 inches, respectively.

In operation, the aircraft aims the weapon in a conventional manner by either moving the launch tube relative to the aircraft to aim it at the target or by aiming the entire aircraft and launch tube at the target. When it is desired to launch the projectile 1, a conventional electrically-ignited squib is activated to ignite the solid propellant spin motors 17. The tangentially directed thrust forces produced by the propellant gases issuing from the nozzles 35 of the spin motors cause the projectile 1 to rotate about its longitudinal axis. A low-friction spin support is provided at the forward end of the projectile by the fin ring bearing assembly 43, whereby the lin ring 39 and the fins 41 remain relatively stationary while the projectile body rotates therein, and at the rearward end of the projectile by the longitudinally extending skirt portion 29 riding on the bearings 53. After reaching its designed spin rate of, for example, 200 revolutions per second, the boost motor propellant grain 11 is electrically ignited to accelerate the projectile in a longitudinal direction. Since the distance between the trailing edge of the fins 41 and the forward exit end of the launch tube 47 is equal to the distance between the rearward end of the longitudinally extending skirt 29 of the projectile 1 and the forward end of the projectile supports 49, fins 41 pass the forward end of the launch tube at the same time that the rearward end of the projectile passes the forward end of the supports 49 and moves into the larger internal diameter portion of the launch tube. This condition is illustrated in FIG. 5. The projectile is thus in the free flight phase of launch while it is partially contained within the launcher envelope. Since the fore and aft projectile supports simultaneously release from the launch tube, and since the clearance in the forward section of the launch tube is sufficient to prevent projectile contact under substantially all launch conditions, tip-off errors normally associated with tube launched projectiles are eliminated.

As the missile proceeds along its fiight path toward its intended target, it is stabilized aerodynamically by the rotationally free fin ring 19 at its forward end, which tends to remain stationary because of the aerodynamic forces exerted upon it by the moving air stream, and the flared skirt 21 at its rearward end. The projectile is also spin stabilized to reducethetendency of the missile longitudinal axis to rotate as aerodynamic torques are introduced around the missile center of mass. The aerodynamic torques occur when an angle of attack disturbance causes a net lifting force on the missile body. The point of application of the net force lis known as the aerodynamic center of pressure and this point is, in general, not coincident with the missile center of mass. The relative separation distance between the center of pressure and center of mass combined with the net lifting force gives rise to a torque about the mass center which tends to rotate the body axis in a direction which is dependent upon whether ythe force is located fore or aft of the mass center. Besides providing spin stabilization, the spin imparted to the missile by the spin motors 17 also reduces ight path deviation caused by radial center of mass and/ or thrust asymmetries. This combination of aerodynamic and spin stabilization offers a gross improvement in delivery accuracy over that obtained with either aerodynamically stabilized or spin stabilized conventional unguided rocket weapons. Since tip-off is eliminated and the rocket supports, i.e., fins 41 andskirt 21, remain as an integral,`v aerodynamically useful, part of the projectile during its flight, high delivery accuracy is achieved without `producing the debris normally associated with tube launched projectiles.

In view of the above, it will be seen that the several objects of the invention are achieved and other advantagebus results obtained.

AS various changes could lbe made in the above constructions without departing from the scope of the invention, it is intended that all matter contained in the above description or shown in the accompanying drawings shall be interpreted as illustrative and not in a limiting sense.

What is claimed is:

1. A iin and spin stabilized projectile, which comprises: t

an elongated cylindrical body having a forward and a rearward portion, said rearward portion having a larger circumference than said forward portion for substantially the length of said rearward portion along a part comprising an integrally constructed, flared, aerodynamic skirt, said skirt comprising a first projectile support during launch;

propulsion means carried by said rearward portion for propelling said projectile along a flight path to a target;

means, carried by said projectile, for rotating said projectile about its longitudinal axis for spin stabilizing said projectile during fiight; and

a plurality of symmetrically arranged and rotationally free aerodynamic lin means carried by said projectile for guiding said projectile along a flight path to a target, said -fin means comprising a second projectile support d-uring launch.

2. A projectile as set forth in claim 1 wherein said plurality of fin means is carried by said forward body portion and constitutes a forward projectile support.

3. A projectile as set forth in claim 2 wherein said plurality of fin means comprises a ring surrounding said projectile body with four aerodynamic fins spaced 90 de- :grecs about said ring and extending radially outwardly therefrom, said ring being journaled to said body for rotation relative thereto.

4. A projectile as set forth in claim 3 wherein said ared skirt comprises an angularly flaring portion and a longitudinally extending portion, the latter constituting said rearward support.

5. A projectile as set forth in claim 1 wherein said pro-` pulsion means comprises a solid propellant rocket motor and said means for rotating comprises a solid propellant spin motor having a generally tangentially directed exhaust nozzle.

6. A combination iin and spin stabilized projectile and launcher therefor, which comprises:

an elongated projectile body having a for-ward and a rearward portion, said rearward portion comprising an integrally constructed, flared, aerodynamic skirt, said skirt comprising a first projectile support during launch;

propulsion means carried by said rearward portion for propelling said projectile along a tight path to a target; v

means for rotating said projectile about its longitudinal axis for spin stabilizing said projectile during ight;

a plurality of symmetrically arranged and rotationally free aerodynamic iin means carried by said projectile forvguiding said projectile along a ight path to a target, said tin means comprising a second projectile support during launch;

an open ended projectile bearing tube; and

a plurality of supports spaced about the inner periphery of said tube so as to define first and second tube sections delineated by first and second internal diameters such that said first projectile support rests in said first internal diameter and said second projectile support rests in said second internal diameter.

7. The combination as set forth in claim 6 wherein the internal diameters of said first and second tube sections are substantially equal to the diameters of said rearward portion and said iin means of said projectile, respectively.

8. The combination as set forth in claim 7 wherein said plurality of spaced supports each comprise a longitudinally elongated retainer secured to the inner wall of said tube, and bearings carried by said retainer for providing a low-friction longitudinal and rotational projectile support.

9. The combination as set forth in claim 8 wherein the distance between the rearward edge of said fin means and the forward end of said tube is equal to the distance between the rearward end of the projectile skirt and the forward end of said longitudinally extending supports.

10. A rocket system, which comprises:

a smooth bore, projectile bearing tube;

a plurality of supports spaced about an inner periphery of and attached to said tube, said supports and smooth bore tube defining -irst and second tube sections delineated by first and second internal diameters;

a projectile carried by said tube prior to launch;

a first self-contained projectile support consisting of a portion of projectile body, said body comprising an elongated cylindrical structure with forward and rearward integral portions, said rearward portion having a larger circumference than said forward portion for substantially the length of said rearward portion, said rearward portion comprising said first projectile support;

a second self-contained projectile support comprising a plurality of symmetrically arranged and rotationally free aerodynamic tins carried by said forward portion of said projectile body, said projectile adapted to be carried by said tube such that said first projectile support rests in said first tube internal diameter and said second projectile support rests in said second tube internal diameter.

11. The rocket system of claim 10 wherein said plurality of spaced supports individually comprises a longitudinally elongated retainer secured to the inner wall of said tube along said lfirst tube section, and bearings carried by said retainer for providing a low friction longitudinal and rotational projecting support.

12. The rocket system as set forth in claim 10 wherein the internal diameters of said first and second tube sections are substantially equal to the diameters of said rearward portions and said aerodynamic projectile lins, re-

spectively.

13. The rocket system as set forth in claim 12 wherein the distance between the rearward edge of said projectile tins and the forward end of said tube is equal to the dis- 7 8 tance between the rearward end of said projectile skirt 3,338,134 8/ 1967 Barnes 89-1.808 and the forward end of said longitudinally extending sup- 3,251,267 5/1966 Hauser et al. 89-11808 ports. f "3,427,921 2`/19 69 Bode et al. 89-1.808

References Cited UNITED STATES PATENTS 2,968,996 1/1961 Strickland et al. 89-1.808 U.S. C1. `X.R.

3,084,600 4/1963 walker 102- 93 f 3,225,655 12/1965 Inglis :x9- 1.816 4276R891816'14R244-323 5 SAMUEL W. ENGLE, Primary Examiner UNITED STATES PATENT @ENCE CERTIFICATE OF CORRECTION Dated QQLQDQ: 5 j 91] Patent No. 3 .610 .O96

Inventor(s) Rober@ It is certified that error appears in the above-identified patent and that said Letters Patent are hereby corrected as shown below:

Signed and sealed this 21 st day of March 1972.

(SEAL) Attest:

EDWARD M.FLETCHER,JR. ROBERT GO'ITSCHALK Attestng Officer' Commissioner' of Patents FORM *D0-05D (10'59) uscoMM-nc como-peo 7 U S` GOVERNMENT PRINTING OF'ICE: IPI 0-355334