US3188958A - Range control for a ballistic missile - Google Patents

Description

June 15, 1965 J, D. BURKE ETAL 3,188,958

RANGE CONTROL FOR A BALLISTIC MISSILE Filed March 11, 1963 3 Sheets-Sheet 1 Efi N 5%; w me m 0: ug 2 a E 3 E O O 8 I James D. Burke Robert J. Parks :2 Robert M. Stewart INVENTORS. 1!. 4.0? A9 Mai/9 Mud June 15, 1965 J. D. BURKE ETAL RANGE CONTROL FOR A BALLISTIC MISSILE 3 Sheets-Sheet 2 Filed March 11, 1963 James D. Burke Robert J. Parks Robert M. Stewart,

INVENTORS.

FIG. 4

RANGE CONTROL FOR A BALLISTIC MlSSlLE James D. Burke, Altadena, Robert J. Parks, La anada,

and Robert M. Stewart, Eucino, Califi, assignors, by

mesne assignments, to the United States of America as represented by the Secretary of the Army Filed Mar. 11, 1963, Ser. No. 265,589 8 Claims. (Cl. 102-50) This invention relates to a range control system for ballistic missiles and, more particularly, to a highly accurate dragbrake range control system for surface-tosurface missiles.

The success of surface-to-surface weapon systems depends on a reliable and accurate method for controlling range. There are several basic considerations involved in the design of a range-control system for any surfaceto-surface missile.

One major problem in the development of ballistic missiles has been finding a suitable range guidance system for substantially eliminating target error. This may be solved by using either a thrust-termination system or a dragbrake system. Although the addition of dragbrakes to a missile presents serious aerodynamic, structural, and mechanical design problems, it is believed that these problems can be solved with less difliculty than would be required for a thrust-termination system.

A problem in the development of a highly accurate ballistic missile using dragbrakes is the type of propellant system to be used. A liquid propellant system has many advantages. Liquid propellant missiles have a flightdemonstrated record for accurate impulse control. The manufacture of components and the production of liquid propellants involve standard industrial capabilities. On the other hand, a solid propellant rocket motor having the same range capability could be designed with significantly less weight and len th. The solid propellant rocket motors consist basically of only four components: the motor case, the nozzle assembly, the propellant grain, and the igniter. Thus, the number of components attributable to a solid propellant motor is substantially less than the number required for a liquid system. This reduction in components reduces complexity, and thus a solid propellant system has a greater inherent reliability. In addition, the use of solid propellants considerably simplifies the supply system. There is no separate element in the supply line that is associated with the propulsion system. The propellant travels as part of the missile. In contrast, the supply of oxidizer, fuel, and spare propulsion components for a liquid system requires complex storage, handling, and shipping facilities from the point of manufacture to the battle front. Finally, because the propellant is already loaded in the solid rocket, it places no demand on the tactical organization. This organization, when handling liquid propellants, is forced to form a continuation of the supply line. Therefore, considerable checkout equipment and a large number of highly trained personnel must be devoted to testing and fueling before the liquid propulsion system is in a condition comparable to that of the solid rocket as delivered.

Another basic consideration involved in the design of a range control system for a ballistic missile is the type of guidance system to be used. One type of guidance system which integrates well with both warhead requirements and the general characteristics of a ballistic missile is the so-called vernier type of control. A vernier type system is one in which all effects that could cause dispersion from the target are cancelled, either as early as they can be detected or as shortly thereafter as is practical. In a rocket-powered missile, essentially all of the large dispersion effects occur during the burning period which is a relatively small fraction of the total flight time.

areas If a coarse-guidance system which operates either during burning or shortly thereafter is used to overcome these large dispersion effects, any further guidance need be only of a fine or vernier nature. This vernier guidance needs to have only the capability of overcoming the relatively small disturbances which occur later in flight in orderto provide the final accuracy.

Thus, the range-control problem has two separate aspects, coarse and fine control. Coarse control implies the means by which gros variations are effected between minimum and maximum range; all settings of the guidance equipment for the coarse range selection are known and made before firing, and all of the major control effects must take place early in flight. Fine control implies compensation during flight for unpredictable variations in motor performance, drag coeflicient, missile weight or alignment, winds, air density, and temperature. This compensation must be made after burnout and afterthe period of high drag on the ascending phase. Thus, by overcoming the major dispersion efiects early in flight and designing the fine guidance system with only a limited capability, the period when system failures can cause large target dispersion can be limited to a relatively short period early in flight.

Another consideration concerns the type of maneuverability that should be used to provide the primary adjustment and control of range for a surface-to-sur'face ballistic missile. It is a characteristic of the ballistic type of trajectories to be used by such surface-to-surface missiles that during and shortly after the burning period the range is highly sensitive to axial acceleration, whereas late in flight it is sensitive to pitch acceleration and quite insensitive to axial acceleration. The use of axial maneuverability allows the major disturbances (which oc-' our early in flight and are primarily axial-acceleration disturbances) to be corrected shortly after they appear and permits the vernier-system concept to be used. It also allows the missile to be designed with very low lateralmaneuvering capabilities so that, duringthe last part of the flight, forall practicalpurposes, nothing (not even system failures) can cause the missile to deviate significantly from its prescribed trajectory.

A final consideration is the type of guidance to be used with the dragbrake system of this invention. The guidance system must incorporatesorne type of trajectorymeasuring equipment to provide information for the actuation of the dragbrakes for range guidance. Such measurements could be made by either air-borne, allinertial elements, or by ground-based radio system or by a combined radio-inertial system. The present invention utilizes a combined radio-inertial guidance system.

By properly combining the inertial data with the radio data, a continuous, non-distorted, up to the minute estimate of the missile position and velocitycan be made with an accuracy significantly better than that attainable by either system alone. In order to obtain high accuracy from radio data, it is necessary to filter the data over a relatively long period. This filtering normally introduces time-lag or distortion errors; however, the inertial data can be used to eliminate these errors and still allow the long filter time. The stabilized inertial data also allows radio guidance to be terminated near the peak of the tra jectory, thus essentially eliminating any radio horizon considerations, while still providing for the accurate measurement and correction of trajectory disturbances which occur after the fine-dragbrake period.

One object of this invention is to provide an improved dragbrake system for accurately controlling the range of a ballistic missile.

Another object of this invention is to utilize a solid propellant rocket motor in conjunction with a dragbrake system for controlling the range of a missile.

Patented June 15, T1965 A further object of this invention is to use a Vernier type of guidance control for a surfaw-to-surface ballistic missile to provide greater target accuracy.

Another object of this invention is to provide a missile which has low maneuverability during flight.

It is a further object of this invention to use a combined radio-inertial guidance for controlling the flight path of a surface-to-surface missile.

According to the present invention the foregoing and other objects are attained by providing an inertially guided missile with an improved dragbrake range-control system. The dragbrake system comprises a single mechanism having four paddle-like brake members which open and close for a coarse and fine braking period. During the coarse brake period the brakes are opened in response to a pre-set timing mechanism and are closed by a signal from a coarse brake closing computer. During the fine brake period-the brakes are opened by a signal from a fine brake opening computer and closed by a signal from a fine brake closing computer. a

The invention will be more fully understood through the following detailed description taken in conjunction with the accompanying drawings wherein like reference numerals designate identical or corresponding parts throughout the several views, in which:

FIGURE 1 is a schematic view of a ballistic missile embodying the invention;

FIGURE 2 is a view along lines 2-2 of FIGURE 1;

FIGURE 3 is a view of the dragbrakes in the closed position showing the hydraulic actuating cylinder; 7

FIGURE 4 is a partial View of the dragbrakes in the extended position;

FIGURE 5 is a view of the gimbal system used in the inertial guidance system; and I FIGURE 6 is a schematic block diagram ofthe range control system of the present invention.

Referring to FIGURE 1 a missile 1th is shown having a forward section 12, a central section 14, and an aft section 16. The forward section contains a warhead 13 and is a separate unit which is attached by conventional means shortly before the missile is to be fired. Central section 14 contains the inertial guidance equipment 17, and an autopilot 18 which actuates control surfaces 28 and jet vanes in order to roll-stabilize the missile and'to yaw and pitch the missile as required. A dragbrake actuator system 29 controls the operation of fourwedge-shaped dragbrakes 22. Dragbrakes 22 are located immediately aft of the guidancerequipment and a solid propellant rocket motor 24 is located directly aft of the dragbrakes. The missile is stabilized by four fixed deltashaped fins 26 which have movable plain flat-type trailingedge control surfaces 28 attached thereto. let vanes 30 are mechanically interconnected with and used in com bination with control surfaces 28 to provide initial stability and to give reduction in maneuverability, after burnout, that is desired for reasons of warhead safety.

Referring to FIGURE 2 the four dragbrakes are shown in fully extended position. Dragbr-akes 22 which are identical in size and shape operate in unison and are paddle-like members each having a six-sided periphery such that all sides are unequal and unlike in contour. As.

shown more clearly by FIGURE 4, side 52 is rounded to conform with the outer skin'of the missile. The end surface 34 lies at'an angle of approximately 90 to surface 32 and is rounded so that, in closing, the brake has has a snug fit. Side 36 is a short straight side positioned at about a angle from side 34. Side 38, which lies at about an angle of 45 q with respect to side 36, is also straight and serves to stop the dragbrakes when they are in the closed or retracted position. .Sides .4!) and '4-2 are straight and form a combined line which connects with side 32. The six sides of each dragbrake form a rough triangle, the apex of which contains an integrally mounted pivot rod 44 which is pivotally secured to the airframe near the periphery of the missile. A short actuating or linkage arm 46 has one end pivotally connected about rod 44 and the other end pivotally connected to another rod 47. A second actuating or linkage arm 48 is pivotally connected to rod 47 and a third rod 49 which is mounted near the end of one of the arms of a sprocket 5t}. Sprocket 50 is in turn free to rotate about a central hub 51 which is axially aligned with the longitudinal axis of the missile and secured in a suitable manner to the airframe of the missile. A pneumatic cylinder 52 is also suitably mounted to the airframe of the, missile and is used to rotate sprocket 59 which controls the operation of the dragbrakes. A piston 54, which is movably mounted within cylinder 52, has a connecting rod 56 extending out one end of the cylinder and attached to the hub of sprocket 5t). Piston 54 is operated by a gas or fluid medium contained in a dragbrake actuating system 26 which is fed to cylinder 52 through conduits 58 and 59. When the dragbrakes are in the closed position (see FIGURE 3) the pressure is exerted through conduit 59. When the dragbrakes are in the extended position as indicated by FIGURE 4, the pressure is exerted through conduit 58.

' FIGURE 5 schematically indicates that a portion of the inertial guidance system used in this missile includes a stabilized platform 60 which is mounted in the missile and stabilized relative to space to establish a space fixedg missile-guidance system of three reference coordinates. The inner axis is chosen as pitch axis 62, the middle axis as yaw axis 64, and the outer axis as roll axis 66. The platform may be defined as a device to place three accelerometers to predetermined earth-referenced space coordinates, and to maintain this orientation during the flight of the missile. The accelerometers used are the force-balance type with a feedback network such that the amplifier output voltage contains components proportional to the azimuth acceleration, velocity and displacement. Three gyroscopes are also mounted on the platform to prevent movement of the accelerometers from the established reference axis. The output of the stable platform system consists of three acceleration signals and three angle signals. One accelerometer measures the acceleration perpendicular to the azimuth plane; A sec ond accelerometer measures the acceleration in the azimuth plane perpendicular to the standard trajectory impact tangent. The third accelerometer measures the acceleration along the standard impact tangent. The angle signals indicate the angle between the missile airframe and the pitch, yaw and roll gimbals of the platform.

The ground to air system for controlling the actuation of the dragbrakes is shown in FIGURE 6. The ground system includes radio communications equipment 7 1 for tracking the position of said missile and providing measurements of range, range rate, and elevation. These measurements are sent to coarse-brake closing computer 72, fine-brake closing computer 73, and fine-brake opening computer '74. Each of these computers will required four to eight pre-set parameters. All of theseparameters are related, however, and are primarily a function of target range.

The coarse-and-fine-brake closing computers could be physically the same unit if provision were made for switching of these parameters sometimes between the coarseand-fine dragbrake'periods. These computers are shown separately on the block diagrams for clarity. Fine-brake closing computer 73 and fine-brake opening computer 74 must be capable of operating simultaneously.

The coarse-brake period is initiated by an adjustable timer 89, the setting being a function of the standard target range.

timer 89 to actuate relay 87. Relay 87 in turn causes dragbrake actuating system 20 to exert pressure through conduit 58 (see FIGURE 4) to extend the dragbrakes. As the missile approaches the correct time for coarsebrake closure, coarse brake-closing computer 72 automa- At the preset time a signal is sent from tically begins to operate due to the appropriate range signal fed from radio equipment 71 and the computer output continuously estimates how far beyond the target the missile would impact if the brakes were closed at that particular moment. This output is sent through switch 76 to radio communication equipment 71 Where it is transmitted to airborne communications system 80. Here it is filtered by low-pass filter 81 of a particular bandwidth and mixed in mixer 84 with acceleration data from inertial guidance system 82. When the value of this combined signal goes through zero the brakes are closed by actuation of a dragbrake closing relay 85. When the dragbrakes close, switch 77 is closed by a suitable relay means (not shown). The output of fine-brake opening computer 74 continually estimates the time before the standard time (referenced to impact) at which the brakes will actually close if they are opened at that particular time and extended long enough to give the correct range. This signal is sent to airborne communications system 80 Where it is filtered by another low pass filter 83 of a different bandwidth and mixed at mixer 86 with acceleration data from inertial guidance system 82. When the value of this combined signal goes through zero the dragbrake opening relay 87 is actuated to cause the dragbrakes to be extended. When the brakes open switch 76 is actuated by a suitable relay means (not shown) to connect fine brake closing computer 73 with radio communications 71. The output of computer 73 continually estimates how far beyond the target the missile would impact if the brakes were closed at that particular moment. Again, this signal is received by airborne communications system 80 where it is filtered and mixed at point 84 with acceleration data from inertial guidance system 82. When the value of this combined signal goes through zero the brakes are closed for the final time by actuation of relay 85.

The dragbrake operation thus consists of a coarse and a fine braking period. The desired range determines when these braking periods occur. For example, on shortrange flights the brakes are initially opened by timer 89 during the early part of the rocket motor burning period. When maximum range is desired, the brakes are initially opened shortly after the end of the rocket motor burning period. The brakes are closed at a time computed to correct for any variations from standard performance up to that point. As missile nears the peak of its trajectory, the brakes are again opened on command and are similarly closed when the drag has again compensated for nonstandard performance. After the missile reaches the peak of its trajectory, the remaining flight is controlled by inertial guidance. The missile has very little lateral maneuver capability. Thus, when the dragbrakes close for the last time, no disturbance (or failure) that subsequently occurs can radically change the point of impact.

It is to be understood that the forms of the invention that are herein shown and described are preferred embodiments, and that various changes are possible in the light of the above teachings. For example, a plurality of braking periods might be employed instead of two braking periods. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described.

What we claim is:

1. In a missile having an airframe, a range control system comprising:

(a) a dragbrake assembly positioned within said airframe near the central portion of said missile,

(b) said dragbrake assembly including a plurality of individual dragbrakes rotatably attached to said airframe;

(c) said airframe having a plurality of spaced openings in the periphery thereof, said openings being disposed radially outwardly of and adjacent to said dragbrakes;

(d) means for initially extending said dragbrakes through said openings in a direction normal to the longitudinal axis of said missile at a predetermined time; s

'5 (e) first means for retracting said dragbrakes within said airframe to complete a coarse dragbrake period of operation which improves range accuracy by eliminating major trajectory disturbances;

(f) second means for again extending said dragbrakes through said openings;

(g) thirdmeans for finally retracting said dragbrakes within said airframe to complete a fine dragbrake period of operation which further improves range accuracy by eliminating minor trajectory disturbances; and,

(h) means for sequential actuation of each of said first, second, and third extending and retracting means.

2. A range control system as set forth in claim 1 wherein:

(a) each of said individual dragbrakes is a wedgeshaped member having a plurality of unequal sides designed to provide aerodynamic stability to said missile during operation of said dragbrakes.

3. A range control system as set forth in claim 2 wherein:

(a) said dragbrake assembly includes a sprocket means disposed for connection to said individual dragbrakes through linkage arms;

(b) pneumatic means connected to said sprocket means to provide a bi-directional rotation to said sprocket means,

(c) rotation of said sprocket means in one direction causing said dragbrakes to extend outside said airframe, and rotation in the opposite direction causing retraction of said dragbrakes within said airframe.

4. A range control system as set forth in claim 3 wherein:

(a) said means for initially extending said dragbrakes includes an actuating mechanism which controls the operation of said pneumatic means,

(b) said mechanism responsive to the operation of a dragbrake opening relay, and

(c) said relay being responsive to the operation of a timing device which is set prior to the launching of said missile.

5. In a missile having an airframe, a range control system comprising:

(a) a dragbrake assembly positioned within said airframe near the central portion of said missile,

(b) said dragbrake assembly including a plurality of individual dragbrakes rotatably attached to said airframe,

(c) said airframe having a plurality of openings near 55 The central portion of said missile for permitting opening and closing of said dragbrakes;

(d) actuating means for extending and retracting said dragbrakes;

(e) tracking means for providing data indicating the range of said missile;

(f) computer means on the ground responsive to said data for determining the correct time for extending and retracting said dragbrakes,

(g) said computer means generating an output signal;

(h) inertial guidance means for providing data indicating the acceleration of said rocket;

(i) means for combining said output signal with said acceleration data to provide a resultant signal; and

(j) relay means responsive to said resultant signal for operation of said actuating means in a series of sequential coarse and fine adjustments of said dragbrakes.

6. A range control system as set forth in claim 5 75 wherein said dragbrakes are initially extended in response r v 7 to a signal from a timing device which is set prior to the launching of said missile. V

7. A range control system as set forth in claim 6 wherein said computer means includes a firstcomputer for generating a signal to retract said dragbrakes to complete a coarse dragbrake period which improves range accuracy by eliminating major'trajectory disturbances,

3.'A range control systemas set forth in claim 7 wherein: p I

(a) said computer means includes a second and third computer, i (b) said second computergenerating'a signal for extending said dragbrakes to initiate a fine dragbrake period of operation, and (c) said third computer generating a signal for finally retracting said dragbrakes after the minor trajectory disturbances have been eliminated.

7 References Cited by the Examiner UNITED STATES PATENTS SAMUEL FEINBERG, Primary Examiner.

Claims (1)

1. IN A MISSILE HAVING AN AIRFRAME, A RANGE CONTROL SYSTEM COMPRISING: (A) A DRAGBRAKE ASSEMBLY POSITIONED WITHIN SAID AIRFRAME NEAR THE CENTRAL PORTION OF SAID MISSILE, (B) SAID DRAGBRAKE ASSEMBLY INCLUDING A PLURALITY OF INDIVIDUAL DRAGBRAKE ROTATABLY ATTACHED TO SAID AIRFRAME; (C) SAID AIRFRAME HAVING A PLURALITY OF SPACED OPENINGS IN THE PERIPHERY THEREOF, SAID OPENINGS BEING DISPOSED RADIALLY OUTWARDLY OF AND ADJACENT TO SAID BRAKE DRAGBRAKES; (D) MEANS FOR INITIALLY EXTENDING SAID DRAGBRAKES THROUGH SAID OPENINGS IN A DIRECTION NORMAL TO THE LONGITUDINAL AXIS OF SAID MISSILE AT A PREDETERMINED TIME; (E) FIRST MEANS FOR RETRACTING SAID DRAGBRAKES WITHIN SAID AIRFRAME TO COMPLETE A COARSE DRAGBRAKE PERIOD OF OPERATION WHICH IMPROVES RANGE ACCURACY BY ELIMINATING MAJOR TRAJECTORY DISTURBANCES; (F) SECOND MEANS FOR AGAIN EXTENDING SAID DRAGBRAKES THROUGH SAID OPENINGS; (G) THIRD MEANS FOR FINALLY RETRACTING SAID DRAGBRAKES WITHIN SAID AIRFRAME TO COMPLETE A FINE DRAGBRAKE PERIOD OF OPERATION WHICH FURTHER IMPROVES RANGE ACCURACY BY ELIMINATING MINOR TRAJECTORY DISTURBANCES; AND, (H) MEANS FOR SEQUENTIAL ACTUATION OF EACH OF SAID FIRST, SECOND, AND THIRD EXTENDING AND RETRACTING MEANS.

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