US3153367A

US3153367A - Anti-missile system

Info

Publication number: US3153367A
Application number: US829146A
Authority: US
Inventors: Frederick W Ross; Kaswen Morris
Original assignee: Individual
Current assignee: Individual
Priority date: 1959-07-23
Filing date: 1959-07-23
Publication date: 1964-10-20
Anticipated expiration: 1981-10-20

Description

Oct. 20, 1964 F. w. R055 ETAL ANTI-MISSILE SYSTEM 2 Sheets-Sheet 1 Filed July 23, 1959 IIIIIIIIIIIIIIIIIIIIIIIIIIIIII! I'll.

ATTORNEYS 1964 F. w. R055 ETAL 3 53,367

ANTI-MISSILE SYSTEM 2 Sheets-Sheet 2 Filed July 23. 1959 United States Patent 3,153,367 ANTI-MESSHLE SYSTEM Frederick W. Ross, Del Mar, and Morris Kaswen, San

Diego, Calif assignors, by mesne assignments, to the United States of America as represented by the Secretary of the Army Filed .iuly 23, 1959, Ser. No. 829,146 4 Claims. (Cl. 89-1) The present invention relates generally to ballistic missile weapons, and specifically to a novel weapons system designed for use particularly in regions above the appreciable atmosphere where aerodynamic control is negligible. The invention provides a weapon of offense and defense. In the following description, it will be discussed primarily as a defensive system designed to intercept and destroy oncoming intercontinental ballastic missiles (ICBM) and satellites which will be referred to herein as targets.

The intercontinental ballistic missile carrying a nuclear warhead is concededly the potentially most dangerous implement of warfare available in any presently known weapons system complex. With a range of excess of 5000 miles and a velocity of more than 10,000 miles per hour, the ICBM is capable of delivering its destructive payload from a distant launching site in a matter of minutes. It was early recognized that a primary problem in anti ICBM defense resides in lack of a reasonable warning time. Without adequate advance notice, the problem of interception practically defies solution. Recognition of the need for early warning resulted in the development of long range radar systems, high speed digital computers and associated components capable of detecting, locating, and evaluating an ICBM threat while the enemy missile still is perhaps several thousand miles from its intended target, thus allowing a defense command from five to ten minutes to meet and destroy the threat well away from the target area.

The anti-ICBM program has followed a logical development based on the concept of launching a high altitude, high speed missile designed to intercept the impending threat at some distant point in its trajectory where detonation of the ICMB will result in no deleterious effects to the intended target. However, it has been established that such long range, high altitude interceptions necessitate requirements which exceed the capabilities of existing ground based guidance systems. In other words, given the benefit of all necessary data regarding the enemy ICBM, ground control cannot consistently or even generally come within a lO-degree initial error range in launching a defensive missile Weapon.

Midcourse and terminal controls have thus been utilized in an effort to correct for initial error. Midcourse control, actuated either from the ground or from missile borne instruments, has been found to be unable to correct the defensive missile path to within effective distance of the planned intercept point.

Thus, the burden has fallen on terminal control guidance of the defensive weapon to effectuate a hit. Prior to the present invention, terminal guidance has consisted solely of side-thrust maneuvering, which may be generally summarized as follows:

Extremely high closing velocities between the enemy target missile and the defensive weapon limit available control time to a matter of seconds even with reasonably long range missile borne seekers or fire control systems, resulting in demands for unattainably high g load maneuvers even for small initial errors against nonmaneuvering targets. Extremely high target speeds make impractical the design of a defensive interceptor weapon with a speed advantage. This severely restricts permissible flight geometry at the inception of terminal guidance.

3,153,367 Patented Oct. 20, 1964 The lack of an aerodynamic medium to provide support for control surfaces on the defensive weapon shifts the burden of control completely to reaction thrusting devices, resulting in need for complex rocket chamber geometry and serious fuel weight penalties.

As an example of the foregoing problems: assume an enemy ICBM in free flight approaching at 16,000 feet per second; assume a defensive missile launched with only a 10-degree initial error in the direction of interceptor weapon velocity, which may be about 4000 feet per second; and assume further a target tracking device operative along a line of sight having a range of 80,000 feet. Although the assumed figures are favorable to the intercepting weapon missile, an error of approximately 3000 feet will develop and must be overcome by terminal control within four seconds if a hit is to be scored, thus making necessary a 10 g side thrust applied without delay.

It is apparent that the possibility of achieving a hit under the assumed favorable conditions is practically nonexistent. Further, it is obvious that if the target were a satellite with a minimum velocity on the order of 26,000 feet per second, the above problems would be of significantly greater magnitude.

An object of the present invention is to provide a new means and method of defense against ICBM and satellite weapons based on utilization of presently existing ballistic missiles and hardware, but with a novel terminal subsystem containing equipment'capable of accurately aiming and projecting an explosive charge, as for example an atomic shell, from the defense missile during the terminal phase.

Another object of this invention is to provide a combination of a ballistic missile and an explosive charge or war head-projecting weapon for improved defense against high velocity objects from ground level to altitudes in excess of 50 miles.

Still a further object of this invention is to provide a novel terminal stage sub-system for use in an anti-ballistic missile weapon system. Such system involves means to place a gun in close proximity to an object at high altitude, to stabilize such gun in relation to the object, and to fire such gun so as to destroy the enemy oncoming missile with the explosive charge or war head projected from the gun.

Other objects and advantages will become more fully apparent from the claims, and from the following description when taken in conjunction with the annexed drawings in which:

FIGURE 1 is a side elevation view in section of the terminal stage of the ballistic missile in accordance with the present invention;

FIGURE 2 is a partial plan view taken along lines 2-2 of FIGURE 1 illustrating certain details of construction in the elevational control for the gun;

FIGURE 3 is a partial elevation view of the fire control system taken along lines 33 of FIGURE 1;

FIGURE 4 is a bottom view of FIGURE 1 showing the disposition of stabilizing exhaust or rocket chambers;

FIGURE 5 is a block diagram of the control system; and

FIGURE 6 is a diagrammatic illustration of the operation of the present invention as applied to the destruction of an oncoming target missile or satellite.

Referring now to FIGURE 1, the terminal stage 10 of the missile in accordance with the present invention comprises a support 12 on which the azimuth ring gear 14 is mounted for rotation by motor 16. Gun barrel 18 is supported on a pair of stanchions 20 for elevation adjustment by mot0r22. Pinion 24 on motor 22 drives gear rack 26 shown also in FIGURE 2 to control the elevation of gun barrel 18 in an obvious manner.

Gun barrel 18 may have a diameter of substantial size to be adapted for firing nuclear explosive type shells and is thus adapted to support the directional antenna 28 which is mounted to be adjustably positioned both in azimuth and in elevation by

motors

30 and 32 respectively. As will become apparent from below, this type of mounting simplifies aiming of the gun, though it should be understood that the antenna may be mounted on other portions of the missile if desired.

Antenna 28 serves as a sensing element for a fire control system which is capable of first seeking the oncoming target missile and then looking on to follow the target missile. Such fire control systems which may be either of the radar or the infra-red type are well known in the art and typically comprise a system as is generally illustrated in the block diagram form in FIGURE 5.

Missile It is provided with nose cone 34 which either folds away or is otherwise removed once the missile rises above the atmosphere. In known anti-missile defense missiles where extreme g load side thrust is required in the terminal stage or phase, multi-stage missiles are practically a requisite, each stage dropping off at burnout so the final or terminal stage will represent the least possible weight factor. It has been considered necessary to separate the terminal stage from the initial and mid-course stages in order to reduce the mass of the missile in the final or terminal stage to the greatest possible degree to maximize the available maneuverability.

However, in the present invention the final stage is essentially a ballistic rather than thrust stage, and it may prove desirable to retain the pre-terminal stage as the frame or base from which to operate the recoilless gun 18. This may effect an overall reduction in the initial weight and complexities necessitated by the jettisoning mechanisms for the various stages while providing increased mass for the terminal stage frame to thereby make the gun aiming apparatus more efficient.

This does not preclude use of separable propulsion units for multiple stages, except that the base vehicle during the terminal phase of the controlled flight must have a moment of inertia appreciably in excess of that of the gun. The ability to rotate the gun, which if nuclear explosives are used, must necessarily be massive, through an angle in relation to the carrying vehicle depends on the reaction moments supplied by the base. Thus, the gun itself may be angularly accelerated relative to space by means of reaction thrusting devices produced by combustion products from rocket chambers 33 oriented as shown in FIGURE 4 or by universally mounted exhaust tubes 35 and 37 shown in FIGURE 1 or by a combination thereof. The total angular motion of the base vehicle during gun sluing and tracking will be approximately that of the gun decreased by the ratio of gun inertia to platform inertia.

Further, the actual attitude of the platform is only of secondary importance since the gun is referenced directly to the target by line-of-sight sensors, for example of the infra-red type. The primary vehicle needs no close stabilization of its motions particularly where antenna 28 is mounted to the barrel of gun 18. So long as the angular position of missile 16 is maintained within reasonable limits, e.g. to degrees about some design value, the fire control problem is not adversely affected. It will be appreciated that the tracking system antenna 28 need not be directly attached to the recoilless gun 18 as shown in FIGURE 1, but may occupy any suitable space in missile 10.

Referring now to FIGURE 5, a block diagram of the fire control system is shown which includes the antenna 28 and transmitter-receiver 40 arranged to pick up the target when line-of-sight is established. The tracking system, controlled by the servos and 32, picks up the oncoming target missile or satellite and transmits data to the computer 44, which may be of a conventional type that evaluates stored data and the data newly supplied by the receiver part of transmitter-receiver 40, and

provides a signal which is determinate of the angular displacement necessary for the recoilless gun 18 to project its War head to a point to assure collision. Computer 44 directs the

gun servo motors

16 and 22 which aim the gun, set the time fuse if desired, and fire the war head upon command. Primary power for the tracking system, computer, and gun servo motors is supplied by power source 46.

Summarizing, in the present invention the primary ballistic missile 10 is thus used merely as a carrier of a suitable recoilless gun 12 and a fire control system of the homing type which acquires or picks up and tracks an oncoming target missile or satellite as soon as the line of sight contact is established. Tracking signals are sent to computer 44 which determines and transmits control signals to elevation and

azimuth motors

16 and 22 for gun 18. Gun 18 is so aimed that the gun line is along a line which provides a collision course for the guns projectile. It is contemplated that the war head will achieve an actual collision as a general rule. However, the shell if of atomic capability will be capable of destroying an ICBM or satellite at a radius of several hundred feet or much more, depending on the shells capacity, and a time fuse may be included in the shell to activate the War head, in which case the computer provides a time fuse signal. Since the predicted shell flight time is accurately known under the existing near vacuum conditions, correct computation is a relatively simple task.

FEGURE 6 illustrates the basic geometry of operation of the invention to achieve interception and destruction of the oncoming target 48 where a prior art missile would register a complete miss. With reference to FIGURE 6, assume that target 48 has been picked up by early warning radar system and determined to be following a trajectory above the atmosphere which will take it along path A-B at a velocity of 16,000 feet per second.

Assume a prior art missile 50 with a velocity of 4000 feet per second fired along path CD and designed to intercept the target 28 above the atmosphere but not equipped with the present invention. Assume further a nominal initial course error of only 10 degrees for missile it Under these conditions, missile It} will be at point Y when the target 48 crosses the path 0-1) of missile 50 at Z and a clean miss of some 3000 feet will result.

However, by utilizing the present invention in missile in with the same assumptions regarding the target and the same initial course error of missile it), the following sequence will ensue:

As the missile 1i clears the atmosphere, the nose cone 34 (FIGURE 1) folds away or is otherwise disposed of, and the fire control system antenna 23 begins scanning. As the missile 1% and the target 48 come Within about 80,000 feet of one another, the fire control system establishes a line-of-sight detectionand locks ontotarget 43. From data determined by the relative movement of the tracking system as it maintains direct bearing on target 48 and predetermined data, computer 44 determines the direction and magnitude of error in the course of missile 10, and directs positioning and firing of the recoilless gun 18 along a path as indicated by the guns projectile velocity vector such that the war head is projected along path F-G to intercept and destroy the target 48 at point X. The war head fired by gun 18 may be armed with a time fuse which may be set to explode the war head at the determined point of closest proximity of warhead and target should it be determined by the computer 44 that there exists possibility that the war head will not eifect actual collision.

The system, operating in the unique environment atforded by near vacuum conditions present above the atmosphere, ofiers significant advantages in prediction capabilities and projectile range. The ability to operate in a free fall condition for both target and interceptor, since neither is maneuvering, introduces a unique form of straight line trajectories which permits considerable simplification of the involved computations and instrumentation required in known fire control systems.

The defensive Weapon provided by the present invention is capable of greatly increased maneuverability, is insensitive to short closing times, leads to a substantial weight decrease from known defensive missile systems, has a remarkably high intercept accuracy, and reduces the requisite accuracy required of the ground control system.

It can be readily appreciated that by providing the flexible, easily maneuvera'ole terminal arrangement above disclosed, the invention substantially reduces and minimizes the necessity for extremely accurate launch and midcourse control. In effect, with the invention, the launching phase need only direct the missile with reasonable accuracy, and midcourse control is required only to bring the terminal system to a point where a line-of-sight is established.

It is obvious that many modifications may be made to the basic system as disclosed, and the foregoing embodiment is to be considered in all respects as illustrative and not restrictive, the scope of the invention being indicated by the appended claims rather than by the foregoing description, and all changes which fall within the meaning and range of equivalency of the claims are therefore intended to be embraced therein.

What is claimed and desired to be secured by United States Letters Patent is:

1. In a ballistic missile having a charge for destroying a target moving in a ba listic trajectory outside the earths atmosphere, an apparatus for directing said charge into a collision course with said target comprising: propulsion means located in the aft end of said missile for lifting said missile out of the earths atmosphere; auxiliary rocket stabilizer motor means attached to said missile; a recoilless gun mounted in the forward end of said missile; scanning means in the forward end of said missile for detection of a target above the earths atmosphere; means in the'forward end of said missile for projecting the path of said target object detected by said scanning means and aligning said gun, relative to a projected path of the missile, with a point in the projected path of said target object; means for firing said auxiliary rocket stabilizer motors to stabilize said missile during alignment of said gun; means for firing said charge from said gun; and means removably enclosing said forward end of said missile.

2. In a ballistic missile having a charge for destroying a target moving in a ballistic trajectory outside the earths atmosphere, an apparatus for directing said charge into a collision course with said target comprising: a rocket motor positioned in the aft end of said missile for lifting said missile out of the earths atmosphere; auxiliary rocket stabilizer motor means attached to said missile; a recoilless gun movably mounted in the forwardend of said missile; said charge being carried by said gun; scanning means in the forward end of said missile fixedly attached to said gun for line-of-sight detection of a target object above the earths atmosphere; means in said forward end of said missile responsive to signals from said scanning means for projecting the path of said target object and aligning said gun, relative to a projected path of the missile, in azimuth and elevation with a point in the projected path of said target object; means for selectively firing said auxiliary rocket stabilizer motor means to stabilize said missile during alignment of said gun; means for firing said charge from said gun into a collision course with said target object; and a nose cone removably enclosing said forward end of said missile.

3. A ballistic missile as claimed in claim 2 wherein said auxiliary rocket stabilizer motor means is comprised of a plurality of selectively fireable rocket motors fixedly attached to the aft end of said missile in a plane perpendicular to the longitudinal axis of said missile.

4. A ballistic missile as claimed in claim 2 wherein said auxiliary rocket stabilizer motor means is comprised of a plurality of selectively fireable gimballed rocket motors.

References Cited in the file of this patent UNITED STATES PATENTS 1,103,503 Goddard July 14, 1914 1,294,240 Cooke Feb. 11, 1919 2,395,435 Thompson et a1 Feb. 26, 1946 2,399,426 Bradley Apr. 30, 1946 2,410,723 Edwards et a1 Nov. 5, 1946 2,433,843 Hammond et al Jan. 6, 1948 2,933,980 Moore et al Apr. 26, 1960 2,938,459 McGraw et a1 May 31, 1960 OTHER REFERENCES Astronautica Acta, vol. 3, No. 1, 1957 (pages 1-15), The Uses of Artificial Satellite Vehicles (Part II) by Canney and Ordway (pages 3-5, The Artificial Satellite V elliole as a Military Weapon).

Claims

1. IN A BALLISTIC MISSILE HAVING A CHARGE FOR DESTROYING A TARGET MOVING IN A BALLISTIC TRAJECTORY OUTSIDE THE EARTH''S ATMOSPHERE, AN APPARATUS FOR DIRECTING SAID CHARGE INTO A COLLISION COURSE WITH SAID TARGET COMPRISING: PROPULSION MEANS LOCATED IN THE AFT END OF SAID MISSILE FOR LIFTING SAID MISSILE OUT OF THE EARTH''S ATMOSPHERE; AUXILIARY ROCKET STABILIZER MOTOR MEANS ATTACHED TO SAID MISSILE; A RECOILLESS GUN MOUNTED IN THE FORWARD END OF SAID MISSILE; SCANNING MEANS IN THE FORWARD END OF SAID MISSILE FOR DETECTION OF A TARGET ABOVE THE EARTH''S ATMOSPHERE; MEANS IN THE FORWARD END OF SAID MISSILE FOR PROJECTING THE PATH OF SAID TARGET OBJECT DETECTED BY SAID SCANNING MEANS AND ALIGNING SAID GUN, RELATIVE TO A PROJECTED PATH OF THE MISSILE, WITH A POINT IN THE PROJECTED PATH OF SAID TARGET OBJECT; MEANS FOR FIRING SAID AUXILIARY ROCKET STABILIZER MOTORS TO STABILIZE SAID MISSILE DURING ALIGNMENT OF SAID GUN; MEANS FOR FIRING SAID CHARGE FROM SAID GUN; AND MEANS REMOVABLY ENCLOSING SAID FORWARD END OF SAID MISSILE.