US3411736A

US3411736A - Missile guidance system

Info

Publication number: US3411736A
Application number: US513392A
Authority: US
Inventors: Joseph A Kelly
Original assignee: Motorola Inc
Current assignee: Motorola Solutions Inc
Priority date: 1965-12-13
Filing date: 1965-12-13
Publication date: 1968-11-19
Anticipated expiration: 1985-11-19

Description

Nov. 19, 1968 J. A. KELLY MISSILE GUlDANCE SYSTEM 4 Sheets-Sheet 1 ZINVENTOR-a ose BY j 1" J fi wy, M 3% Filed Dec. 15, 1965 Nov. 19, 1968 J. A. KELLY 3,411,736

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Nov. 19, 1968 J. A. KELLY 3,411,736

MISSILE GUIDANCE SYSTEM Fild Dec. 15, 1965 4 Sheets-Sheet 5 i MISS/LE E 70 r -PILOT Rang;- LI

0 2o 40 so TIME MILL/SECONDS g TOTAL MISSILE 80057 THRUST g PILOT M/ssu FIXED aoosr r/musr 2 5 1g .STAIRCASE RISE TIME TIME AFTER PILOT ROCKET SEPARATION MILL/SECONDS INVENTOR. z9se z .fl BY fla w sew.

Nov. 19, 1968 J. A. KELLY MISSILE GUIDANCE SYSTEM 4 Sheets-Shget 4 Filed Dec. 15, 1965 y 2 w w J M w W a a H United States Patent 3,411,736 MISSYLE GUIDANCE SYSTEM Joseph A. Kelly, Riverside, Calif., assignor to Motorola, Inc., Franklin Park, IlL, a corporation of Illinois Filed Dec. 13, 1965, Ser. No. 513,392 6 Claims. (Cl. 2443.15)

ABSTRACT OF THE DISCLOSURE A rotating pilot rocket is propelled into a chamber within a missile to be guided. The chamber has detection means sensing the location of the pilot rocket within the chamber. Control signals are generated in response to sensed pilot rocket location and are used to guide the missile. The pilot rocket is impelled into the chamber immediately prior to missile launch. Wedge shaped photo Sensors are used in the chamber to sense rocket locations.

This invention relates generally to missile guidance systems, and more particularly to systems for guiding a surface-to-surface missile with respect to a reference trajectory which is generated and contained within the missile itself.

Missiles of the surface-to-surface type can be guided over a ballistic trajectory by providing a proof mass having a reference vector within the missile to retain the missile on a predetermined trajectory. Displacement of the missile with respect to the proof mass provides a control signal to cause corrective action so that the airframe of the missile is maintained on the desired trajectory.

Guidance systems are known for providing a programmed course, either prerecorded to be carried by the missile or supplied to it by radio command, for all or part of the missile trajectory. Inertial guidance systems for all-the-way guidance to the target utilize complex and expensive precision gyro stabilized references to maintain the missile in a desired trajectory, while radio controlled systems are subject to jamming and countermeasure techniques. Further, support of the proof mass within the missile gives rise to torques which cause the reference vector to precess from the initial setting.

A missile guidance system employing a free-fall proof mass to impart a true vacuum trajectory to the missile once the initial boost velocity vector has been established is set forth in copending application Ser. No. 840,781, filed Sept. 17, 1959, now Patent 3,233,848 and commonly assigned.

It is, therefore, an object of the present invention to provide a simple and improved inertial guidance system for missiles.

Another object is to provide a method of total guidance for a missile so that the missile will follow a predetermined reference vector during initial boost and a precise vacuum trajectory that is free from atmospheric disturbances subsequent to the initial boost.

A further object of the invention is to provide a guidance system for missiles wherein a simple self-contained control system provides complete control of the missile so that it follows the magnitude and direction of a predetermined boost velocity vector.

Still another object is to provide a guidance system for missiles wherein a self-contained control system provides Patented Nov. 19, 1968 control of the missile with respect to a self-propelled proof mass to allow the missile to follow a predetermined boost trajectory and a true vacuum ballistic trajectory subsequent to the boost interval.

A feature of the invention is the provision of an inertial guidance system including a self-propelled proof mass, and displacement sensors to produce control signals indicative of the relative position of the missile with respect to the proof mass to guide the missile so that it follows the trajectory imparted to the proof mass.

A further feature is the provision of a self-propelled pilot rocket, rotating at high speed and contained within a missile to provide a velocity vector for the missile, with displacement sensors to provide variable thrust control of the missile in response to the relative position of the pilot rocket in the chamber in the missile. Lateral displacement signals command the missile to follow the direction of travel of the pilot rocket and longitudinal displacement signals match the missile acceleration and velocity to that of the pilot rocket.

A still further feature is the provision of a self-propelled rotating proof mass contained within a missile to produce a spin axis to be followed by a missile, with displacement sensors having outputs causing control of the boost engines of the missile in response to the relative position of the missile with respect to the proof mass, so that e velocity and direction of the spin axis provided by the proof mass are followed by the missile. When the propellant for the proof mass is expended, the proof mass is retained in a free-fall condition within the missile to provide a ballistic trajectory to which the trajectory of the missile is matched to provide guidance of the missile for the entire flight.

The invention is illustrated in the drawings wherein:

FIG. 1 diagrammatically illustrates a missile having a guidance system in accordance with the invention;

FIG. 2 diagrammatically shows the self-propelled reference mass and the sensing chamber in accordance with this invention;

FIG. 3 diagrammatically shows the device for sensing the position of the reference mass in the chamber in a combined quasisectional-perspective diagrammatic view used to clearly illustrate the relationships of the component parts;

FIG. 4 is a block diagram illustrating the control system;

FIGS. 5 and 6 illustrate modes of operation of the missile system;

FIG. 7 is an enlarged diagrammatic partially sectioned elevational view of the rocket mechanism of FIG. 2;

FIG. 8 is a diagrammatic plan view of the rocket support mechanism of FIG. 2 with a cutaway portion to illustrate some constructional features.

In practicing the present invention there is provided a guidance system for a self-propelled missile of the ballistic type capable of a high-g inclined boost trajectory. The missile includes engines having fixed boost thrust and variable sustainer thrust capabilities, and means for lateral control so that the acceleration and direction of the missile can be controlled with respect to a proof mass during the entire flight. The proof mass is a pilot rocket which is launched into an enclosed chamber within the missile and remains unrestrained by and unattached to the missile so as to provide a self-propelled torque-free reference vector to be followed by the missile during the boost period. Shortly after launching of the pilot rocket the missile engines are ignited to provide a thrust which rises to a level matching the acceleration of the pilot rocket. Displacement sensors located around the periphery of the chamber detect the relative position of the missile with respect to the pilot rocket to provide signals for control of magnitude and direction of the thrust imparted to the missile so that it follows the magnitude and direction of the reference vector provided by the pilot rocket. The pilot rocket is calibrated for thrust and aligned for direction prior to launch to set an accurate course to be followed by the missile and when launched is rotated at high speeds to provide a spin axis which resists precession by external forces. This spin axis is fixed in space to provide the direction of boost for the missile so that the missile follows a predetermined boost trajectory.

At the end of a calibrated impulse period the propellant for the pilot rocket is expended and it continues in free fall inside the chamber in the missile. At this time the velocity of the missile is identical (within small tolerances) with that of the pilot rocket and missile boost thrust is terminated. A variable sustaining thrust for the missile is maintained to overcome drag, and control over both lateral movement and acceleration of the missile is retained by sensing displacement between the free fall trajectory of the pilot rocket and the missile to provide guidance for the remainder of the trajectory.

Referring now to the drawings, in FIG. 1 a ballistic type missile incorporating the guidance system of the invention is shown generally at in a launch position preferably at 45 or higher from the horizontal. The missile has an intermediate forward portion 11 adapted to carry a warhead or a pay load, and a nose portion 12 carrying an autopilot system shown generally at 13. This autopilot system includes angle-of-attack transducer 15 and control unit 16 containing a yaw rate sensor and associated electronic circuitry for the autopilot system. A power pack 17 contains batteries and necessary power supplies for operation of the autopilot system. The autopilot system forms no part of the present invention and will not be further described.

Rearward portions of the missile include a pair of canted boost, variable thrust liquid engines 20, liquid fuel tanks 22, and servo controlled valves 23 for controlling the flow of fuel to engines 20. Each engine has a fixed canted nozzle 26 and a bob-Weight 27 containing damp ing fluid for roll rate stabilization of the missile. Valves 23 of known design have a fixed area orifice for providing a fixed initial thrust to the missile and a variable area orifice for controlling acceleration during boost and providing a sustaining thrust subsequent to termination of the main boost thrust. The output of autopilot system 13 includes signals to control the valves 23 in response to the position of the missile with respect to a proof mass, all in a known manner. For the canted engines shown, mounted on a missile adapted to spin at a stabilized roll rate, sympathetic control of fuel to the engines provides acceleration control of the missile, while differential control of fuel to the engines provides lateral control of the missile. It is to be understood, however, that the autopilot system 13 is equally adaptable to control aerodynamic lift and drag surfaces as well as other types of variable thrust engines. It is possible, for example, to utilize a single canted, variable thrust engine with a single fuel control value for providing lateral and speed control of the missile in response to the output of theautopilot system, or to control canard surfaces of aerodynamically controlled missiles propelled by ram-jet engines. The environment in which the guidance system may be practiced is not limited to the illustrated missile system.

The precise guidance system of the missile includes a self-propelled proof mass, or pilot rocket 30, shown in its prelaunched position as lodged in the conical socket of retaining assembly 32, as shown. Pilot rocket 30 is adapted to be projected into guidance chamber 28, located near the missile center of rotation, immediately prior to launch of the missile. The relative position of pilot rocket 30 with respect to the center of chamber 28 is detected by displacement sensors to provide information signals for control of both acceleration and direction of the missile.

As shown in detail in FIG. 2, wherein the proof mass or pilot rocket 30 is illustrated in both its prelaunch and post launch positions, and as seen in more detail in FIGS. 7 and 8, pilot rocket 30 is held for prelaunch alignment by conical retaining assembly 32 on tripod assembly 37. Bearing assembly 38 rotatively supports retaining assembly 32. As shown, retaining assembly 32 may be integrally formed on the outer race of bearing assembly 38. The conical socket of retaining assembly 32 releasably holds pilot rocket 30 permitting rotation thereof with respect to the airframe. A jet of high pressure gas injected through a hollow one of the support arms of tripod assembly 37 from gas source 39 and directed into the air pockets 33 of bearing assembly 38 imparts a spin to the pilot rocket 30 while held by the retaining assembly 32 in the illustrated prelaunch position. As best seen in FIGS. 7 and 8, the high pressure gas flows through a conduit to the inner race of bearing assembly 38 and is directed against air pockets 33 formed in the inner race of bearing assembly 38. Typically, this spin may be in the order of revolutions per second.

The front face 31 of pilot rocket 30 is highly polished and is used to align the pilot rocket spin axis in the azimuth plane to the target at the proper elevation angle. By use of a simple optical projection system of the periscope type including projection tube 42, inserted in the forward end of chamber 28, an image is projected to and reflected from the polished front face 31 of pilot rocket 30. The projected and reflected images are then compared in a ground coordinate system (not shown) such as by known optical superposition techniques, to enable alignment corrections of the missile azimuth and elevation angles in its launch carriage to insure precise alignment of the spin axis of the pilot rocket with a distant target on the basis of a ballistic trajectory. Since the orientation of the spin axis of the pilot rocket provides the reference vector for the trajectory, alignment may be made without regard for the missile orientation. Alignment in this manner provides accuracies better than one minute of arc in both azimuth and elevation angles.

Ignition circuit 44 and ignitor 45 ignite the solid propellant charge 46 of pilot rocket 30* to cause it to move forward into chamber 28 away from retaining assembly 32. The ignition gasses of pilot rocket 30 are exhausted through gas duct 47 by the exhaust action of the main missile engines. The thrust of the propellant varies with temperatures, therefore, heater 48 blows a stream of hot air over pilot rocket 30 prior to launch to maintain its temperature constant to enable a precisely calibrated impulse to be imparted to it. For a given rocket, at a constant temperature the calibrated impulse is determined by the amount of the propellant and its burn rate. Immediately after ignition of the pilot rocket solid propellant charge, the main boost engines of the missile are ignited so that both the missile and the pilot rocket move forward with a predetermined boost acceleration. Maximum missile thrust is relatively greater than the rocket thrust to enable the missile to maintain its accelerationsuch as to position rocket 30 centrally of chamber 28. Once injected into chamber 28 the position of pilot rocket 30 with respect to the center of the chamber is detected by displacement sensors 40 disposed on the walls of chamber 34.

The pilot rocket displacement sensing system may be of the type shown in FIG. 3, wherein pilot rocket 30, in its post-launch position, is viewed from the forward end of chamber 28. For non-aerodynamic missiles of the type shown in FIG. 1, provided with a relatively constant roll rate, it is only necessary to sense displacement of the missile with respect to the pilot rocket along two axes: one parallel to the missile axis for thrust and drag matching; and one at right angles to the missile axis, rotated with the missile about the spin axis of the pilot rocket, for lateral control. Sensing in a single plane allows control in all lateral directions and the outputs of the displacement sensor of FIG. 3, derived along these two named axes, provide signals for polar control of the constantly rotating missile.

The displacement sensor of FIG. 3 uses a ball-shadow technique to detect the position of pilot rocket 30 relative to the walls of chamber 28. A plurality of alternating wedge-shaped selenium photocells 50 are used to detect the position of shadow 52 of pilot rocket 30. The wide tnds of each row of photocell wedges are connected to a differencing amplifier to detect any displacement of the shadow with respect to the center of the chamber. The wedge-shaped photocells are disposed on two walls of the chamber, with the longitudinal axes of the wedges on each wall perpendicular to and parallel to, respectively, the spin axis of pilot rocket 30 to provide error signals relative to the center of the chamber in both the longitudinal and lateral direction.

To form a shadow that is independent of the distance between pilot rocket 30 and photocells 50, a luminescent panel 54 is provided on the opposite two walls of chamber 28. Honeycomb arrays 56, lined with light absorbing material, collimate the light from panels 54. Battery operated power supplies 58 provide the energy necessary for light panels 54.

With reference to FIG. 4, the longitudinal output signal of the displacement sensor is coupled to autopilot system 13 through electrical filter 61 of known electrical design. This signal is unidirectional, having an amplitude indicative of the relative fore-aft position of pilot rocket 30 with respect to the center of chamber 28. Autopilot system 13 has two

channels

13a and 13b, one to control each of the servo valves 23 for engines 20. The longitudinal output signal of the displacement sensor provides command signals of the same polarity to each channel to provide a control signal for values 23 so that for any given instant in time there is sympathetic control of the flow of fuel to the engines to result in a simultaneous increase or decrease of the speed of the missile to match the speed of the pilot rocket.

Since the missile is spinning at a predetermined roll rate, the output signal of the lateral displacement sensor is sinusoidal at the roll frequency of the missile. Typically this is in the order of cycles per second. The output of the angle of attack transducer and the yaw rate sensor, also sinusoidal and at the roll rate frequency, are further coupled to the servo amplifiers of autopilot system 13. Proper phasing is accomplished by positioning the sensors relative to the plane of the canted nozzles of engines 20.

As shown in FIG. 4, the instantaneous lateral output of the displacement sensor, as well as the outputs of the yaw rate sensor and the angle of attack transducer, supply signals to each of

channels

13a and 13b of the autopilot system which are 180 out of phase. Thus a differential control signal, phased with the plane of the nozzles of engines 20, are applied to the servo controlled valves 23 of each engine. This produces differential thrust modulation of the missile engines, synchronized with its roll rate, to result in lateral control of the missile. The yaw rate sensor and the angle of attack transducer provide a damping effect to substantially reduce the deviation of the missile about its center of gravity in the presence of sudden disturbing forces, while the lateral output of the displacement sensor provides the precise lateral control of the missile with respect to the reference vector provided by pilot rocket 30.

In operation, after alignment to the target in the azimuth plane and at proper elevation angle by optical alignment device 42, pilot rocket 30 is spun at about 100 revolutions per second by high pressure gas driving conical socket assembly 38. See FIGS. 7 and 8 wherein air from a leg 37 is directed against air-pocket 33 formed in assembly 38, as shown. The solid fuel propellant of the pilot rocket is then ignited to cause it to move forward into sensor chamber 28 at a slow rate. Shortly after pilot rocket 30 leaves socket assembly 38, the missile booster engines are ignited to give the missile an initial preset thrust for matching the initial acceleration of the pilot rocket. Superimposed on this preset thrust is a variable thrust which is servo controlled in the above described manner to provide speed control for the missile and to null any off-center relative position of pilot rocket 30 with respect to chamber 28. The missile follows the pilot rocket until bumout, at which time the fixed boost thrust is terminated. The variable thrust, however, is retained as a sustainer thrust to counteract drag on the missile and to provide the necessary speed and lateral control. Subsequent to burn out, pilot rocket 30 continues in free fall inside of chamber 28 and continued servo control of the sustainer thrust compensates for missile drag and provides for lateral control with respect to the ballistic trajectory provided by pilot rocket 30.

The magnitude of the boost-end velocity imparted to pilot rocket 30 can be set by calibrating its impulse to mass ratio. Increments of mass such as metal rings 60 can be removed from the pilot rocket to select a boost-end speed for range determination of the missile with extreme accuracy. By way of example, although not limiting, a typical pilot rocket providing a 40 mile ballistic trajectory with a boost-end velocity of 2700 feet per second is about one inch in diameter and weighs about two ounces. Using 'known types of solid propellant, a pilot rocket of this size can conveniently be provided with two pounds of thrust for four seconds and with a total impulse calibrated for better than one part in one thousand.

To insure that the missile is brought up to thrust shortly after the pilot rocket leaves its prelaunch alignment socket and before it has preceded to the forward end of sensor chamber 28, it is necessary that the fixed missile boost have a faster thrust rise time than the pilot rocket. This condition is readily met by use of a known slow burning solid propellant for the pilot rocket, and fast response liquid engines for the missile. Control over the missile can be achieved either by providing it with an extremely fast initial rise time for a predetermined interval, or by providing it with a relatively slower rise time which is servo controlled by the displacement sensor over a some what longer interval. Typical pilot rocket and missile initial boost acceleration curves for the two mentioned conditions are illustrated in FIGS. 5 and 6.

'In FIG. 5 the initial response times of both the missile and the pilot rocket are shown as substantially linear for illustration although in practice they may vary somewhat in linearity for the same results since the thrust is integrated to obtain boost-end velocity. The missile engines, ignited approximately 25 milliseconds after launching of the pilot rocket, are provided with an initial faster response than that of the pilot rocket. This allows the missile to catch up with the pilot rocket so that it is substantially centered in chamber 28. Thus, as shown in FIG. 5, the pilot rocket reaches a constant acceleration of 20 Gs approximately 100 milliseconds after launch. The missile, on the other hand, reaches a constant acceleration of 20 Gs approximately 50 milliseconds after its engines are ignited or milliseconds after launch of the pilot rocket. At that time the velocity of the missile with respect to that of the pilot rocket is well within the variable thrust control capabilities provided by the longitudinal output of the displacement sensor.

For existing known propellants for both the pilot rocket and the missile, the maximum velocity of the pilot rocket relative to the missile, which occurs at crossover point 70 of the two acceleration curves of FIG. 5, is approximately 4 feet per second. By the time the missile and the pilot rocket velocities are equal, point 71 on the curves of FIG. 5, the pilot rocket has moved relative to the missile less than three inches in chamber 28. During this time the longitudinal output of the displacement sensor becomes effective to control the portion of the missile boost engines which can be varied, and the velocity of the missile is matched to that of the pilot rocket for the remainder of the flight trajectory.

For missiles where it is desirable to employ a slower initial boost response, the pilot rocket is provided to have a relatively slower rise time, such as a 450 millisecond rise time to reach a G acceleration as shown in FIG. 6. The fixed portion of the missile boost is provided to have a relatively slow staircase rise time of 420 milliseconds to establish the 20 G acceleration of the missile. This can be readily accomplished by a single solenoid operating a sliding gate valve to provide equal millisecond time delays between successive openings of the fixed boost orifices supplying fuel to the missile engines. The variable thrust portion of the missile engines has a {faster rise time, in the order of 50 milliseconds, and is used to match the missile to the pilot rocket during the initial boost interval. As shown in FIG. 6, the delay time between pilot rocket ignition and missile boost ignition is approximately 100 milliseconds. Shortly after the initial rise time, the variable portion of the missile engine is controlled by the longitudinal output of the displacement sensor to reduce the total effective boost thrust. This acts to maintain the position of the pilot rocket in the center of chamber 28 for the remaining period of the boost period. At the end of approximately 450 milliseconds, both the pilot rocket and the missile have reached a constant 20 G acceleration and matching of the missile velocity with the velocity of the pilot rocket is accomplished in the manner previously discussed.

Although the guidance system has been disclosed for use with variable thrust control of non-aerodynamic missiles, it should be apparent that the guidance system of the invention can be used in missiles having canard control surfaces for aerodynamic control. In addition, other types of engines can be provided to give the missile the desired boost velocity and to provide sustaining thrust for the remainder of the flight.

The guidance system can be further used with other types of missiles rather than the ground-to-ground surface missile shown. For example, the guidance system can be used in a missile launched from an aircraft, with a portion of the initial velocity being imparted to the missile by the aircraft. Since it is the attitude of the pilot rocket rather than the entire missile which determines the trajectory to which the missile follows, it is only necessary that the pilot rocket and not the entire missile be aligned with the target. The missile itself may be carried in any convenient position by the aircraft and subsequent to launch it will automatically align its axis with the reference vector provided by the pilot rocket in the manner previously discussed.

In longer range applications and in application for satellite injection where gravity turn of a missile having a fixed boost engine may be employed, it is further possible to incorporate a throttling capacity into the pilot rocket. This may be accomplished by mounting a small linear accelerometer in the nose of the pilot rocket to measure its accelerations to provide cutoff after a predetermined time. By way of a subminiature radio link located on the pilot rocket the longitudinal output of the displacement sensor can be used to control a variable area orifice valve on the pilot rocket to maintain its fore and aft positions in the chamber so that it follows the acceleration provided by the main boost engines of the missiles. The lateral output of the displacement sensor then commands the missile to follow the direction of travel of the pilot rocket to provide lateral control of the missile in the manner previously discussed.

The invention provides therefore a simple yet extremely accurate inertial guidance system for missiles for all-theway guidance to the target. A self-propelled proof mass, completely contained within the missile to be free from atmospheric disturbances and physical restraint, provides accurate control of the magnitude and direction of the boost velocity vector of the missile. At the end of a predetermined impulse interval this proof mass remains in a free fall condition within the missile to provide a reference for a true vacuum ballistic trajectory, thereby retaining complete and accurate control of the missile during its entire flight period.

I claim:

1. A guidance control system for a propelled missile with an airframe having driving means and means for controlling the direction of said airframe in flight including actuating means responsive to first and second control signals respectively for controlling the velocity and direction of the missile so that the missile follows the trajectory of a reference mass during flight, the improvement including in combination:

means within said airframe forming a chamber,

a reference mass adapted to be impelled into said chamber,

guide means for initially aligning said mass with respect to said chamber for predetermining the trajectory to be imparted to said mass, means to cause said mass to be impelled into said chamber immediately prior to launch of said missile,

sensing means including means establishing a reference point in said chamber for providing control signals indicative of position of said mass with respect to said reference point within said chamber during flight of said missile,

2. The system of claim 1 wherein said mass has a propellant, and ignition means for igniting said propellant to provide said mass with a predetermined impulse causing it to be propelled into said chamber, with said mass remaining in free fall within said chamber at termination of said impulse.

3. The system of claim 1 wherein, said sensing means includes first and second portions positioned along two orthogonal walls of said chamber, a source of light disposed along opposite walls of said chamber to project a shadow of said mass onto said sensing means, means responsive to light on one of said walls for providing said first control signal representative of the position of said mass longitudinally of a predetermined point in said chamher, and means responsive to light on the other of said walls providing a second control signal representative of the position of said mass laterally of said point in said chamber.

4. A guidance system for a missile having an airframe with driving apparatus and a control device for controlling the movement thereof, said guidance system including in combination, means forming a chamber within the airframe, a self-propelled reference mass initially on said airframe aimed into said chamber for imparting a predetermined trajectory thereto, means for initiating propulsion of said mass to thereby free it from said airframe and impel it into said chamber, sensing means responsive to the position of said mass in said chamber for providing con- .trol signals representing the position of said mass with respect to said chamber, and means coupling said sensing means to the control device for controlling the movement of the airframe in response to said control signals so that said mass is positioned centrally of said chamber throughout the movement of said airframe.

5. A guidance system in accordance with claim 4 and including means for aligning the direction of thrust imparted to said mass with respect to said chamber and wherein said sensing means includes light sensitive means positioned along two orthogonal walls in said chamber, light producing means disposed on opposite walls of said chamber to produce a shadow of said mass on said light sensitive means, means responsive to said light sensitive means to provide control signals representative of the position of said mass with respect to a reference point in said chamber, and means for applying said control signals to the control device for controlling the movement of said missile so that said mass is positioned at said reference point throughout the flight of the missile, whereby the trajectory of said missile is determined by the trajectory imparted to said self-propelled reference mass.

6. A guidance system in accordance with claim 4 and including an annular bearing assembly rota-tably supporting said rocket on said airframe, and means to selectively rotate said rocket on said bearing assembly.

References Cited UNITED STATES PATENTS 2,852,208 9/ 1958 Schlesrnan 244-311 3,073,550 1/1963 Young 244-3.l1 3,233,848 2/ 1966 Byrne 244--l4 0 BENJAMIN A. BORCHELT, Primary Examiner.

V. R. PENDERGRASS, Assistant Examiner.