|Publication number||US3374967 A|
|Publication date||Mar 26, 1968|
|Filing date||Dec 6, 1949|
|Priority date||Dec 6, 1949|
|Publication number||US 3374967 A, US 3374967A, US-A-3374967, US3374967 A, US3374967A|
|Inventors||Plumley Harold J|
|Original Assignee||Navy Usa|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (10), Referenced by (41), Classifications (9)|
|External Links: USPTO, USPTO Assignment, Espacenet|
March 26, 1968 H. J. PLUMLEY 3,374,967
COURSECHANGING GUN-LAUNCHED MISSILE Filed DEC. 6. 1949 I 3 Sheets-Sheet l FIRE CONTROL March 26, 1968 H. J. PLUMLEY CHANGING GUN-LAUNCHED MISSILE COURSE- Filed Dec. 6. 1949 5 Sheets-Sheet 7;
O a t March 26, 1968 H. J. PLUMLEY COURSECHANGING GUN-LAUNCHED MISSILE 3 Sheets-Sheet 3 AMPL IF If R Filed Dec.
7'4 R6 E T INTERCEPT/ON RANGE 360 DEGREES 0F RQTAT/OA/ United States Patent I 3,374,967 COURSE-CHANGING GUN-LAUNCHED MISSILE Harold J. Plumley, Washington, D.C., assignor to the United States of America as represented by the Secretary of the Navy Filed Dec. 6, 1949, Ser. No. 131,441 5 Claims. (Cl. 2443.14)
The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.
The present invention relates to apparatus for changing the course of a gun-fired missile during the flight thereof. More particularly, the invention relates to apparatus for changing the course of a rotating, fin-stabilized missile during the flight thereof by firing a reaction steering charge contained in the missile through one side thereof and in a direction at right angles to the line of flight, the steering charge being arranged to direct its force through the center of gravity of the missile.
As a result of the steering force being directed through the center of gravity of the missile, the missile is thrust laterally of the initial trajectory and assumes the course of the new trajecory with its longitudinal axis disposed at an angle thereto, the longitudinal axis of the missile being parallel to the initial trajectory. The stabilizing fins of the missile, however, serve to bring the longitudinal axis into alignment with the new trajectory.
In aiming a missile toward a moving target it is a well known and easily solved problem to compute the proper lead and angle of trajectory when the target is following a known straight line course at a known speed. Where the target changes its course and/or speed after an uncontrolled missile has left the propelling gun, the missile traveling along a set trajectory misses the target and is wasted.
In the use of the present invention, the chance of hitting a target after a change of course thereof, although fortuitous, is enhanced by the firing of the steering charge which alters the course to a predetermined angle such, for example, as five degrees from the original-trajectory. If the course of a missile in flight does not change upon a change upon a change in course of a moving target, the missile will continue along its original set trajectory and would unquestionably miss the target, the kill possibility being nil. In accordance with the arrangement presented by the present invention, the probability of the missile coming in proximity to the target is increased by a ratio of l in 36 over that expected with a fixed trajectory missile.
To achieve this enhanced probability of kill over that expected with fixed trajectory missiles, exhaust ports in communication with an explosive steering charge are provided in the casing of the missile whereby jet action imparts a transverse thrust thereto, while in flight, upon ignition of the charge. Since the invention employs a rotating missile, the transverse thrust may occur anywhere within the 360 rotated by the missile, the angular spin position of the ports at the instant of ignition of a charge being fortuitous in determining the course of the new trajectory taken by the missile. Since the angular orientation of the ports at the instant of ignition of the steering charge is fortuitous, there is a great probability that the ports may be so oriented as to direct the missile in a direction different from that necessary for target interception, resulting in a wasted missile as in a set trajectory missile; on the other hand, there is also the probabilitythat the angular spin position of the ports may be such that the missile will be directed along a trajecory approaching the target. If the ports, at the instant of steering charge ignition, are positioned within a arc, the center ice of which arc is the exact angular spin position for providing a collision course with the target, the trajectories the missile may assume within this 10 are are such as to bring a missile in proximity to the target. Therefore, the probability of the ports being in a favorable angular spin position for the missile to assume a trajectory that approaches the target is 10 in 360, or 1 in 36, or approximately 3 percent.
In the use of the steering charge, it is apparent that on a right angular plane of the original trajectory of the missile at the target interception range there is a gradually decreasing spiral path of points to which the missile may be directed by firing the steering charge during the flight of the missile. If the steering charge is fired shortly after the missile leaves the muzzle of the gun, the corrective distance on the aforementioned plane of target interception at right angles to the original trajectory will be great, the distance becoming increasingly less in the spiral as the firing of the steering charge is delayed during the flight of the missile.
The missile of the present invention is rendered capable of operating in the foregoing manner by the use of the system which will be hereinafter more fully disclosed. The missile is provided with canted stabilizer fins set at an angle suflicient to impart a rotational speed of approximately ten revolutions per second during the flight thereof. The missile is also provided with a modified proximity fuze which has means for transmitting an asymmetrical signal of predetermined frequency rather than the usual symmetrical signal employed to fire the high explosive charge as the missile comes into proximity with a target, the signal being rendered asymmetrical for the purpose of indicating the instantaneous rotational position of the missile in flight and being adapted for reception at the launching point and also reflected from the target according to the conventional operation of the proximity fuze. As is well known to those skilled in the art, the signal reflected from the target and received by the fuze is of a frequency different from the transmitted predetermined frequency by an incremental amount due to the Doppler frequency shift effect. The signal transmitting means of the fuze of the present invention has the multiple functions of transmitting the asymmetrical signal toward the target for reflection back to the fuze, transmitting the signal to the launching point of the missile Where it may be received by suitable apparatus to indicate the rotational position of the missile, receiving a signal from the launching point for firing the steering charge, and receiving the reflected signal thereby to fire the high explosive charge when the reflected signal reaches a predetermined intensity.
An object of the present invention is to provide a new and improved steering control method and system for a gun launched rotating missile having the feature of effectively altering the trajectory of the missile during the flight thereof.
Another object of the present invention is to provide a new and improved method and system of reaction steering control of a missile while in flight.
Another object is to provide a missile having a steering charge mounted therein and capable of producing a known lateral thrust whereby the trajectory thereof may be altered to a predetermined angle at a selected time during the flight of the missile.
Still another object of the invention is to provide a modified firing circuit for a proximity fuze having provision for firing the reaction steering charge when a target-simulated signal is transmitted to the fuze.
A further object is to provide a course changing missile capable of being used with any conventional firing control system to increase the probability of moving target interception above that expected with a fixed trajectory missile.
Still another object is to provide a steering control arrangement for a gun launched missile having suflicient ruggedness to withstand the forces attendant thereto.
A further object is to provide a control arrangement for altering the trajectory of a missile in flight which is compact and of economical construction.
Other objects and many of the attendant advantages of this invention will be readily appreciated as the same becomes better understood by reference to the accompanying drawings wherein:
a FIG. 1 is a view in diagrammatic form of the missile and steering control system of the present invention;
FIG. 2 is a schematic longitudinal sectional view of the missile of the present invention;
FIG. 3 is a sectional view taken along the line 3-3 of FIG. 2;
FIG. 4 is a fragmentary elevation of the tail portion of the missile and illustrating the canted stabilizing fins;
FIG. 5 is a diagram of the electrical system employed in the missile of the present invention;
FIG. 6 is a diagram of the radiation pattern as received at the launching point whereby the rotative position of the missile is made known;
FIG. 7 is a diagram indicating the altered trajectories which may be assumed by the missile during any predetermined one second of the flight of the missile; and
FIG. 8 is a diagram indicating the spiral path corresponding to the trajectories illustrated in FIG. 7.
Referring now to the drawings wherein like reference characters indicate like parts throughout the several views and more particularly referring to FIGS. 2, 3 and 4, the numeral 10 indicates generally a missile constructed in accordance with the present invention. The missile 10 comprises a casing 11 having'a modified proximity fuze 12 mounted in the nose portion thereof, the nose portion being insulated from casing 11 and fuze 12 by suitable dielectric means. At the tail portion of the missile, stabilizing fins 13 are mounted, the fins being canted slightly to rotate the missile at a speed of approximately 10 revolutions per second during the flight thereof. There is mounted in the tail portion a high explosive charge 14. An electroresponsive detonator 15 is provided for firing the charge 14.
Intermediate the nose and tail portions, a steering charge 16 is arranged within the casing 11 and balanced at the center of mass of the missile. The charge 16 comprises a plurality of groups of rocket type explosive discs 17, metallic supporting discs 18 being interposed between the groups of discs 17. In the wall of the casing 11 and adjacent each of the groups of explosive discs 17 there is provided a port 19. If desired, a plurality of the ports 19 may be provided for each of the groups of explosive discs. It is, of course, understood that all of the ports 19 are to be positioned on one side of the casing 11.
Since the missile rotates about 10 r.p.s., the resultant rotation of the angular position of the rocket ports 19 allows the lateral rocket thrust and hence the direction of trajectory change to be made in any direction in the plane normal to the trajectory.
Arranged adjacent the charge 16 at the nose end thereof is an igniter 21 of the electroresponsive type. Bulkheads 22 and 23 separate the compartment containing charge 16 from the tail compartment and the nose compartment, respectively.
The proximity fuze 12 is arranged in the usual manner to ignite the detonator 15 and thereby to fire the high explosive charge when the missile has reached a predetermined distance 'from the target.
In accordance with the invention and as will be more fully described hereinafter, the transmitted radiation pattern of the fuze is asymmetrical in shape and, therefore, it is possible by the use of a suitable receiver at the launching pointto be informed of the exact angular or rotational position of the ports 19 at all times throughout the flight of the missile.
In the diagram of FIG. 5 there is illustrated a modified form of proximity fuze circuit of the Doppler type as employed in the present invention and having a coil 32 mounted transversely in the nose portion of missile 10, the ends thereof being connected respectively to a pair of radiation ears 33 and 34 which are insulatably mounted on the nose portion of the missile to form a transverse dipole. A triode tube 35 has the grid 36 thereofconnected to the coil at one side of the center thereof as at 37 while the coil is grounded to the casing 11 at the opposite side of the center thereof as at 38. Due to the connection of coil 32 to casing 11 at 38, the casing 11 functions as an end-fed dipole and is axially excited by oscillations appearing across coil 32, the frequency of oscillations being determined by the inductance of coil 32 and the interelectrode capacitance of tube 35.
As a result of the arrangement of a pair of dipoles disposed as to radiate fields at right angles to each other and as a result of the antennae connections of coil 32 whereby the dipoles are simultaneously excited, the combined effects of these perpendicularly radiated fields, when viewed by a vertically polarized antenna at the launching point, results in the reception at the launching point of the aforementioned asymmertrical radiation pattern. This asymmetrical radiation pattern may be received at the launching point on .a vertical dipole antenna to give a signal modulated in the form illustrated by the envelope wave 39 of FIG. 6. The asymmetrical envelope wave 39 has high points 41 and lower points 42 which occur when the cars 33 and 34 respectively are facing toward the earth. The steering charge ports 19 are oriented with relation to the ears 33 and 34 in such a manner that, for example, the ports are positioned on the same side of the missile as the ear 33 thereby providing a means of informing the operator at the point of launching of the instant rotational position of the missile at all times throughout the flight thereof. As is apparent from this arrangement and from envelope 39 of FIG. 6, the signal strength which is received at the launching point is proportional to the orientation of the missile or, what amounts to the same thing, is proportional to the angular orientation of the steering ports. When the ports 19 are up, maximum signal strength, as indicated at 41, is received by the receiver; and, when the ports are facing the earth, a signal strength indicated at 42, which is considerably less than the maximum signal strength, is received at the launching position.
The tube 35 is provided with a filament type cathode 43 the circuit thereof comprising a choke 44 to ground on one side and a choke 45 and battery 46 to ground on the other side thereof.
Plate 47 of tube 35 is connected by coupling condenser 54 to an audio amplifier 48, an.A.C. by-pass condenser 51 being interposed between plate 47 and ground and a plate load resistor 52 and B battery 53 also being interposed between plate 47 and ground.
. The output side of amplifier 48 is connected by a C battery 58 to the grid 59 of thyratron 60. Plate 61 of thyratron 60 is connected by a resistor 62 to battery 63 which is connected to ground on the other side thereof. A condenser 64 is connected from the grounded side of battery 63 to the plate 61. The cathode 65 of thyratron 60 is connected to ground through detonator 21 for the steering charge in parallel with an explosive switch 66 which may be of any type suitable for the purpose such, for example, as the explosive switch disclosed in the copending application of Howard C. Filbert, Jr. for Explosive Operated'Pressure Switch, Ser. No. 130,821, filed Dec. 2, 1949, now US. Patent No. 2,721,240 which issued on Oct. 18, 1955. Cathode 65 is likewise connected in open circuit arrangement with detonator 15 for the high explosive charge through the initially open contacts 55 of which 66. As the steering charge is fired, the contacts 55 of explosive switch 66 are closed, thereby connecting the cathode 65 to the detonator 15. It will thus be seen that the first pulsation of tube 60, which pulsation is of short duration, fires the igniter 21 thereby firing the steering charge and simultaneously causing the explosive switch 66 to close the circuit to the detonator 15. The second pulsation of tube 60 fires detonator 15, thereby firing the explosive charge 14.
In the event that the target does not change course and consequently the trajectory of the missile has not been changed, the steering charge will be detonated as soon as the missile is sufiiciently close to the target to render thyratron 60 conductive, and almost instantaneously thereafter thyratron 60 will again be rendered conductive to detonate the explosive charge 14, since the charging time of condenser 64 is very short. In view of the fact that the two conductions of thyratron 60 are succedently instantaneous and in view of the fact that the missile is very close to the target at the time thyratron 60 initially' becomes conductive, it is evident that the missiles course is not sufliciently deflected as to miss the target.
In the operation of the circuit, the output of tube 35, as aforedescribed, produces the asymmetrical wave pattern 39 by reason of the simultaneous radiation from the casing excited axially as a dipole and from the transverse dipole formed by cars 33 and 34.
The receiver 67 at the launching position may pick up the asymmetrical radiation at a predetermined frequency F which is the frequency determined by the inductance of coil 32 and the interelectrode capacitance of tube 35, and may be adapted to feed it to a fire control unit 68, not a part of this invention, which unit, at the prescribed time, as determined by the information retained in the fire control unit, may actuate transmitter 72 at any suitable instant to transmit a target-simulating signal of short duration, the frequency of the target-simulating signal being F -I-AF where AF is a preassigned incremental frequency shift simulating the Doppler effect frequency shift introduced in the signal reflected by the moving target. If no fire control system is used, any suitable manually operated transmitter may be employed to transmit the desired signal. This transmittedsignal modulates the oscillator 35 to derive the difference frequency AF which is applied to audio amplifier 48, amplifier 48 having a narrow pass band centered at AF. Since the limits of velocity range of aircrafts are known and since the velocity of the missile is known, the range of relative velocities between the targets and missile can be calculated, and from these calculations the various Doppler frequency shifts affecting a radiated signal of predetermined frequency and occurring within the range of relative velocities can be determined. From this determination, the pass band of amplifier 48 is designed so as to pass substantially all incremental frequencies caused by the Doppler effect with the aforementioned range of relative velocities, the center frequency of this pass band being AF.
Amplifier 48 amplifies the passed signal AF and ap plies it to the grid 59 of thyratron 60 which is normally maintained at a non-conducting bias by battery 58. The signal transmitted by transmitter 72 is designed to be of sufficient strength to overcome the biasing potential of battery 58 and trigger the thyratron 60.
When the thyratron is rendered conducting, the energy stored in condenser 64 is discharged therethrough by way of plate 61, cathode 65, and thence through the explosive switch 66 in parallel with igniter 21 to ground, thereby to fire the steering charge 16.
Upon operation of switch 66, cathode 65 is in closed circuit arrangement with detonator 15 for the high explosive charge 14. As voltage stored in condenser 64 is only suflicient to fire the explosive switch and igniter 21 and is of short duration, the detonator 15 is not fired during the first signal. As condenser 64 is discharged, the voltage on plate 61 of the thyratron is reduced below the voltage for sustaining conduction and the tube is extinguished. Condenser 64 is charged from battery 63 through resistor 62 which has a resistance sufiicient to reduce the plate potential as aforementioned while also providing for rapid charging of the condenser.
As the missile approaches the target, the radiation pattern of frequency F transmitted by the missile is reflected and shifted in frequency an incremental amount AF by the target, the signal reflected back to the missile assuming the frequency F +AF where AF corresponds to the Doppler incremental frequency shift caused by the velocity of the aircraft. The reflected signal F +AF is mod-ulated in oscillator 35 to derive the difference frequency AF which is passed by amplifier 48 and applied to grid 59 of thyratron 60. The bias 58 is selected to have a value such as to maintain thyratron 60 insensitive to the reflected signals, which increase in intensity as the missile approaches the target, until the missile is within a predetermined distance, for example 40 feet, from the target at which distance the reflected signal is of such intensity as to produce a signal in the output of an amplifier having an amplitude sufiicient to overcome the bias on thyratron 60.
When the amplitude of the output of amplifier 48 in response to the reflected signals reaches a predetermined value to overcome the bias 58 as described in the preceding paragraph, the firing circuit including thyratron 60 operates in the manner heretofore described with the exception that the voltage of condenser 64 is applied to detonator 15, the switch 66 being closed. It will thus be seen that the circuit is arranged in such a manner that on receiving a first signal from the launching point the course of the missile is changed and on receiving a second signal (the reflected signal) the main explosive charge is fired.
Referring particularly to FIG. 1 in which the system for operation of the missile of the present invention is illustrated, the missile 10 is fired from a smooth bore gun 24 along a trajectory A in order to intercept target 25 moving on a course B. Since the projectile travels with a finite speed, the time of projectile flight must be allowed for and the gunis aimed at a point in space ahead of the aircraft on a predicted flight path computed from the targets measured range and vector velocity detected by director 69, resulting in a predicted collision point where course B and trajectory A intersect. At this point of intersection, there is a gradually decreasing spiral path of points in a plane normal to trajectory A to which the missile may be steered, if desired, by firing the steering charge.
Assume that, after firing the missile 10, target 25 changes its course, for example, to course C. In order to change the course of the missile to intercept target 25 at point E, the steering charge 16 of the missile must be fired when the missile reaches the proper rotational position and the proper position along the trajectory to cause such interception by a five degree change of course. Firing of the steering charge is effectuated at random by transmitting from transmitter 72 a signal of short duration which, in turn is received by the missile and fires the steering charge, as aforedescribed. If reception of this signal by the missile occurs when ports 19 are within an arc of 5 degrees on each side of the proper rotational position for attaining course D, then the missile will assume a trajectory approximating course D, as hereinabove described, and eventually comes Within proximity of the target near intercept point E.
Referring to FIGS. 7 and 8, there is illustrated the selective trajectories or altered courses which may occur during one second of the flight of the missile. As will be apparent in FIG. 7, the missile will make 10 revolutions during this period of'time, and the steering charge will be aligned to change the course of the missile in the 0 or 360 position 10 times. The selective trajectories move in a spiral toward the original trajectory, the distance between the convolutions being, in time, one tenth of a second, the forward motion of the missile causing the spiral to become of increasingly smaller diameter as the missile approaches the target interception range. It will, thus be seen that if the steering charge is fired shortly after leaving the gun and at a great distance from the target interception range the spiral will be large and the deviation from the original trajectory great, while if the steering charge is fired at a point near the target interception range, the spiral will be small and the deviation from the original trajectory less. It will, also, be understood that while an angle of five degrees of deviation has been described herein, any desired angle of deviation may be employed by changing the size or force of the steering charge.
It is to be understood that in employing a steering charge of a given force there will be a slight variation in the exact angle of deviation as the forward speed of the missile is reduced during the flight thereof.
It is also to be understood that the present invention is directed to a course-changing missile, per se.
Obviously many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described. I
What is claimed and desired to be secured by Letters Patent of the United States is:
1. Apparatus for changing the course of a missile in flight comprising, a casing for said missile, a reaction charge disposed within said casing and arranged to direct an impulse through one side of said casing and through the center of gravity of the missile as the charge is fired whereby the missile is given a lateral thrust and is directed along a new trajectory with the longitudinal axis of the missile initially at an angle with respect thereto, means on the missile for bringing said axis thereof into alignment with the new trajectory, and means carried by the missile and responsive to a transmitted signal from the launching point for firing said charge.
2. A missile comprising, a casing having a port in one side thereof, a reaction charge mounted within said casing and adjacent said port and arranged to direct an impulse therethrough as the charge is fired whereby the course of the missile is changed, and means including a transmitter located at the launching point of the missile for firing said charge during the flight of the missile.
3. Apparatus for changing the course of a missile in flight comprising, a casing for said missile, canted fins arranged on the tail portion of the missile for maintaining alignment of the longitudinal axis thereof with the trajectory and for causing rotation of the missile about said axis at a predetermined rate, a reaction charge disposed within said casing and arranged to direct an impulse through one side of the casing and through the center of gravity of the missile as the charge is fired whereby the missile is first moved along a new trajectory at an angle thereto and thereafter is brought into alignment there- 8 with by said fins, means in the missile for firing the charge, and means including a transmitter at the launch ing point for igniting said charge firing means while the missile is in flight.
4. Apparatus for performing a course-changing operation on a guided missile subsequent to the launching thereof comprising, a substantially tubular casing for said missile having a plurality of perforations in one side thereof, a reaction steering charge mounted within said casing and arranged to direct a reaction impulse through the perforations thereof as the charge is fired, and means including a transmitter at the launching point for firing said charge during the flight of the missile.
5. In a control system for a gun-launched missile, the combination of an explosive charge disposed within the missile, a reaction steering charge disposed at the center of mass of the missile and arranged to direct a reaction force laterally thereof as the steering charge is fired, canted stabilizing fins on the missile for imparting rotation thereto, a proximity fuze constructed and arranged to transmit an asymmetrical radiation pattern of energy suitable for reflection from the target, and electroexplosive switch means operable concurrently with the firing of the steering charge for connecting the explosive charge in parallel therewith as the steering charge is fired, said fuze having means for firing said steering charge and said electroexplosive switch means in response to energy of predetermined frequency received thereby and transmitted thereto from the launching point, said last named means being arranged to fire said explosive charge in response to said reflected energy as the missile moves into predetermined spaced relation with respect to a target after the steering charge has been fired.
References Cited BENJAMIN A. BORCHELT, Primary Examiner.
JAMES L. BREWRINK, NORMAN H. EVANS,
A. GAUss, J". w. GALLAGHER, H. G. WEISSEN- BERGER, v. R. PENDEGRASS,
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US1288883 *||Jun 24, 1918||Dec 24, 1918||Louis B Harvey||Projectile.|
|US2137598 *||Mar 27, 1936||Nov 22, 1938||Ericsson Telefon Ab L M||Artillery projectile|
|US2145508 *||Dec 27, 1937||Jan 31, 1939||Sageb Sa||Bladed projectile|
|US2403567 *||Jan 13, 1942||Jul 9, 1946||Jr Nathaniel B Wales||Electrically energized fuse|
|US2415348 *||Jun 17, 1943||Feb 4, 1947||John E Haigney||Projectile|
|US2480868 *||Aug 28, 1945||Sep 6, 1949||Charles J Marshall||Guided missile control system|
|US2520433 *||Nov 10, 1941||Aug 29, 1950||Robinson Marion B||Directed missile|
|US2557949 *||Aug 22, 1942||Jun 26, 1951||Standard Telephones Cables Ltd||Fire control system|
|FR832427A *||Title not available|
|IT339479B *||Title not available|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US3695555 *||Jun 12, 1970||Oct 3, 1972||Us Navy||Gun-launched glide vehicle with a mid-course and terminal guidance control system|
|US3758052 *||Jul 9, 1969||Sep 11, 1973||Us Navy||System for accurately increasing the range of gun projectiles|
|US3807274 *||Mar 22, 1973||Apr 30, 1974||Subcom Inc||Method for launching objects from submersibles|
|US3951359 *||Jan 28, 1966||Apr 20, 1976||The United States Of America As Represented By The Secretary Of The Army||Missile control system|
|US3995792 *||Oct 15, 1974||Dec 7, 1976||The United States Of America As Represented By The Secretary Of The Army||Laser missile guidance system|
|US4097007 *||Sep 13, 1976||Jun 27, 1978||The United States Of America As Represented By The Secretary Of The Army||Missile guidance system utilizing polarization|
|US4300736 *||Aug 17, 1979||Nov 17, 1981||Raytheon Company||Fire control system|
|US4533094 *||Aug 23, 1984||Aug 6, 1985||Raytheon Company||Mortar system with improved round|
|US4568040 *||Dec 3, 1982||Feb 4, 1986||Thomson-Brandt||Terminal guidance method and a guided missile operating according to this method|
|US4573648 *||Jan 20, 1983||Mar 4, 1986||Ford Aerospace And Communications Corp.||Ram air combustion steering system for a guided missile|
|US4641801 *||Jan 22, 1986||Feb 10, 1987||Lynch Jr David D||Terminally guided weapon delivery system|
|US4655411 *||Mar 21, 1984||Apr 7, 1987||Ab Bofors||Means for reducing spread of shots in a weapon system|
|US4685639 *||Dec 23, 1985||Aug 11, 1987||Ford Aerospace & Communications Corp.||Pneumatically actuated ram air steering system for a guided missile|
|US4730794 *||May 8, 1987||Mar 15, 1988||Messerschmitt-Bolkow-Blohm Gmbh||Method and apparatus for angle coding|
|US4762293 *||Dec 12, 1968||Aug 9, 1988||The Secretary Of State For Defence In Her Britannic Majesty's Government Of The United Kindgom Of Great Britain And Northern Ireland||Rocket projectiles|
|US4898340 *||Jan 15, 1982||Feb 6, 1990||Raytheon Company||Apparatus and method for controlling a cannon-launched projectile|
|US5071087 *||Mar 11, 1991||Dec 10, 1991||The United States Of America As Represented By The Secretary Of The Navy||Method of guiding an in-flight vehicle to a desired flight path|
|US5238204 *||Jul 28, 1978||Aug 24, 1993||Thomson-Csf||Guided projectile|
|US6231002 *||Mar 12, 1990||May 15, 2001||The Boeing Company||System and method for defending a vehicle|
|US6231003 *||Mar 12, 1990||May 15, 2001||The Boeing Company||Apparatus for defending a vehicle against an approaching threat|
|US6367735 *||Feb 10, 2000||Apr 9, 2002||Quantic Industries, Inc.||Projectile diverter|
|US7004423 *||Jan 29, 2004||Feb 28, 2006||Quantic Industries, Inc.||Projectile diverter|
|US7599814||Apr 27, 2006||Oct 6, 2009||Hrl Laboratories, Llc||System and method for computing reachable areas|
|US7872215 *||Jan 18, 2011||Raytheon Company||Methods and apparatus for guiding a projectile|
|US8058596 *||Nov 15, 2011||Raytheon Company||Method of controlling missile flight using attitude control thrusters|
|US8618455 *||May 27, 2010||Dec 31, 2013||Safariland, Llc||Adjustable range munition|
|US20050103925 *||Jan 29, 2004||May 19, 2005||Mark Folsom||Projectile diverter|
|US20070255524 *||Apr 27, 2006||Nov 1, 2007||Hrl Laboratories. Llc||System and method for computing reachable areas|
|US20100288869 *||Nov 18, 2010||Raytheon Company||Methods and apparatus for guiding a projectile|
|US20110049289 *||Aug 27, 2009||Mar 3, 2011||Kinsey Jr Lloyd E||Method of controlling missile flight using attitude control thrusters|
|US20120175456 *||Jul 12, 2012||Safariland, Llc||Adjustable Range Munition|
|DE2264243A1 *||Dec 30, 1972||Jul 12, 1973||Ship Systems Inc||Lasergesteuertes geschoss|
|DE2500232A1 *||Jan 4, 1975||Jul 15, 1976||Ship Systems Inc||Projectile trajectory correction system by explosives - has longitudinal grooves on projectile circumference charged with high explosive|
|DE2543606A1 *||Sep 30, 1975||Apr 7, 1977||Deutsch Franz Forsch Inst||Verfahren zur steigerung der wirksamen reichweite von geschossen durch impulskorrekturen|
|DE2714688A1 *||Apr 1, 1977||Oct 13, 1977||Bofors Ab||Vorrichtung zur korrektur der flugbahn eines projektils|
|DE4123308A1 *||Jul 13, 1991||Jan 21, 1993||Diehl Gmbh & Co||Guidable projectile for combating re-entry warheads - has hot gas generator drive material, extending as large dia. central core over essential part of projectile length and carrying shrapnel|
|EP0343131A2 *||Apr 28, 1989||Nov 23, 1989||Aktiebolaget Bofors||An apparatus for determining roll position|
|EP0809084A1 *||May 7, 1997||Nov 26, 1997||Tda Armements S.A.S.||Apparatus for determining the roll angle position of a flying device, especially of an ammunition|
|WO1983003894A1 *||Apr 21, 1983||Nov 10, 1983||Hughes Aircraft Company||Terminally guided weapon delivery system|
|WO1984002975A1 *||Jan 20, 1983||Aug 2, 1984||Ford Aerospace & Communication||Ram air combustion steering system for a guided missile|
|WO2001069164A1 *||Feb 7, 2001||Sep 20, 2001||Quantic Industries, Inc.||Improved projectile diverter|
|U.S. Classification||244/3.14, 244/3.22, 244/3.23|
|International Classification||F41G7/20, F41G7/30|
|Cooperative Classification||F41G7/305, F42B10/661|
|European Classification||F42B10/66B, F41G7/30B2|