US3073550A - Guidance system for missiles - Google Patents

Description

Jan. 15, 1963 L. L. YOUNG summers syswm FOR mssms 2 Sheets-Sheet 1 Filed Nov. 4, 1957 INVENTOR.

142m A. Ot/NG Jan. 15,- 1963 L. L. YOUNG 3,073,550

GUIDANCE SYSTEM FOR MISSILES Filed Nov. 4, 1957 FL/Gl/I'CONTBQSYSTEMZ? 2 Sheets-Sheet 2 I INEEUALGU/DANCESKSTEMZ 3 1 INVENTOR. [Ar/22v L. You/v6 United States Patent Office 3,073,550 GUIDANCE SYSTEM FOR Larry L. Young, Downey, Calif.

Filed Nov. 4, 1957, Ser. No. 696,011 12 Claims. (Cl. 24414} This invention relates generally to systems for guiding the flight of ballistic missiles, and more particularly to a missile guidance system which is capable of causing a missile to travel through space as a free mass unaffected by outside influences other than gravity.

It is well known in the science of ballistics that a projectile may be aimed at a target by positioning the launching means at certain azimuth and elevation angles determined by the targets position and the projectiles muzzle velocity. In order to obtain high accuracy at long range targets, corrections to these angles are introduced to adjust the effect of estimated atmospheric conditions along the trajectory, such as barometric pressure, temperature and wind. Such effects are discussed in the book Naval Ordnance and Gunnery, vol. 2, published by the Department of Ordnance and Gunnery, U.S. Naval Academy (1955), page 38 et seq.

With the improvements that have been made in controlling the muzzle velocity of the launching means, and setting the azimuth and elevation angles, the principal remaining cause of inaccuracy is the effect of atmospheric conditions. In a relatively long, high trajectory barometric pressure, temperature, and particularly wind, are quite diflicult, if not impossible, to accurately predict. For this reason, the accuracy capable with unguided projectiles directed at relatively long range targets is seriously limited.

In the newer science of guided missiles, considerable efiort has been expended toward equipping the projectile or missile with some means of guidance during its flight, so that the inaccuracies introduced by these unpredictahle atmospheric conditions can be overcome. In certain applications, guidance systems which control the missile from the ground have been used. Such systems are discussed in a paper entitled Guidance Techniques, presented by Walter L. Webster, Jr., at the Guided Missiles Seminar held by the Advisory Group for Aeronautical Research and Development (AGARD), North Atlantic Treaty Organization in September 1956, and in the book Principles of Guided Missile Design-Guidance, by Arthur S. Locke, published by Van Nostrand Co., Inc., in September 1955, particularly at pages 562 through 567. These ground control systems determine the missiles position by radar or radio, and send control signals to guide it along the path desired. However, such systems have several serious limitations when considered as a means for guiding long range missiles over enemy territory. One is the range limitations on missile-to-ground control communications, and another is the threat of enemy jamming.

Guidance systems have also been developed which are self-contained by the missile and operate on the inertia principle. Such systems are discussed in Electronic Equipment magazine of September 1957 in an article by C. F. ODonnell entitled How Inertial Navigation Works, starting at page 42, and in Missiles and Rockets magazine of April 1957 in an article by J. M. Slater entitled Inertial Guidance for Air Forces High Punch Missiles. The most common of these systems utilizes a gyro stabilized platform and accelerometers all carried within the missile, which measure the accelerating forces to which the missile is subjected. The detected accelerating forces are then integrated against time to give velocity which in turn is again integrated against time to give distance by a computer also carried in the missile, thus 3,073,550 Patented Jan. 15, 1963 providing a continuing indication of the missiles posi tion. The computer then continuously compares the missiles position with a predetermined path and introduces correction to the missiles flight controls to guide it back as nearly as possible to the predetermined path when it deviates therefrom.

The difliculty with such inertia type systems is that .certain inescapable errors, such as acceleration bias and in some cases, as the square or cube of time. These errors must, therefore, be kept extremely small if the missiles guidance is to be accurate. This makes necessary the use of high precision equipment which is quite delicate and subject to frequent malfunctioning. Such equipment is also highly complex and expensive, particularly since it is missile-borne and must therefore meet stringent space, weight and launching shock requirements.

It is therefore a major object of this invention to provide a missile guidance system which is composed of relatively simple components not subject to frequent inaccuracies or failure and yet, due to a new principle of operation, makes possible a very high degree of accuracy.

it is also an important object of this invention to provide a guidance system for missiles in flight which automatically senses and overcomes the effects of unpredictable atmospheric conditions without utilizing missileborne accelerometers or integral computers.

It is another object of my invention to provide a guidance system for missiles which cannot be effectively jammed.

It is a further object of my invention to provide a missile guidance system which operates on the inertia principle, in that after the pro-per initial conditions have been established, the missile is slaved to a free mass which is acted upon only by known forces and therefore follows a predetermined trajectory in accordance with the principles of celestial mechanics.

Still another object of my invention is to provide a missile guidance system of the type described in which a normally free body having mass is fixedly supported with in the missile until the missile has been launched and the proper initial conditions of position and velocity have been established by means of ground control, and is thereafter released to initiate inertial guidance of the missile from within.

Still a further object of this invention is to provide a missile guidance system which can be made small in size and sufiiciently economical to make feasible the guidance of missiles with conventional as well as atomic warheads.

These and other objects and advantages of my invention will become apparent from the following detailed description thereof, read together with the accompanying drawings, in which:

FIGURE 1 is a schematic diagram showing a trajectory typical of a free mass which has a specified velocity, is unaffected by any other force except gravity, and intersects two predetermined points on the earths surface;

FIGURE 2 is a schematic diagram showing ground control means for guiding a missile in the initial stage of its flight;

FIGURE 3 is a perspective view of a missile embodying my invention;

FIGURE 4 is a schematic block diagram of the missileborne controls utilized by my invention;

FIGURE 5 is a perspective view of that portion of the missile which contains the free mass-body that determines the missiles trajectory after its initial stage, showing the mass-body in its fixedly supported condition; and

FIGURES 6 and 7 are perspective views similar to FIG- ity. The eurve 12 represents such an elliptical trajectory beginning at, the point 13 and terminating at the target 11. ,Cur've 12 is aetually one of many curves which could be calculated to extend between point 13 and target 11, but tn keep the velocity requirements low and facilitate guidthe: from the ground during the initial stage of travel,

as ,eirplained later, a curve is chosen which intersects the eart'hs profile about forty-five degrees.

Principle of Operation The principle upon which my invention operates is that if a body having mass is positioned on a free mass trajectory like curve 12 with the proper velocity corresponding to its-particular position, and then, at an instant when these initial conditions are met, completely isolated from all outside influences except gravity (i.e., atmospheric drag, wind, etc.), the body will, by the laws of celestial mechanics, continue along the'trajec tory until it strikes the earths-surfaee at target 11. Since the curve 12 can be calculated quite accurately if the exact coordinates of the target are known, extreme accuracy is possible by utilization of this principle.

It is important to note at this point that velocity, as used herein, means a vector quantity having both speed and direction.

V "In actual operatio'n, I dispose a body having mass in an evacuatedcharnber inside a missile, fixedly support :the body un'til the missile has been launched and the initial conditions of position and velocity established, and then release the body inside the chamber and slave the missile to it by detecting any displacement between the body and the chamber walls and guiding the missile to :m'inimizethisdisplacement. Understanding my general approach, it will be more readily apparent how the various components of my system, hereinafter described, co- :operate to bring about the desired result. Initial Stage Guidance To launch the missile, a launch site 14 is provided ahead of the point 13 under the mathematical trajectory curve 12 and so located with respect to free mass trajectory curve 12 that a conveniently attainable launching trajectory 15 will carry the missile into tangency with curve '12 at point 16 with approximately the necessary velocity.

In order to assure that the missile travels from its launch site 14 to its point of tangencylti and has the proper velocity upon arrival, I provide aground control system 17 The ground control system 17 has three radar rec eivers ISwhich are disposed at different accurately surveyed ground locations near the launch site 14, and

commnnicate means of coaxial cable or line-of-sight ,radio vvith a central control unit 19. 'Ijhe central control "unit 19 has a'radar transmitter Zlland a computer Ell.

The missile itself has a radarrepeaterZZ which retr'an'smits the 'i'adar signals from the transmitterZll. The

re-transmitted signals from the repeater '22 are received "by the "ground-based receivers 18 and c'clmmu'nic'ated to the central control'nnit 19 where they are supplied to the computer 21, along with the signals from'the transmitter. 'Fromthis "information, the computer can continuously calculate the mi's'siles position and velocity.

"l'he coordinates "of the" target and launch site are also suppliedto the computer, and from this information, the

a computer calculates and stores the mathematical equationsiof a family of curves similar to curve 12,. These curves all pass nearthe launch site 14, intersect the earths surface at the target ll, and have a zenith angle 12a close to forty-five degrees; that is, the family of curves includes any 'curve close enough to the bestpossible free mass trajectoryto target 11 .to be effectively utilized as curve 12, and for this reason, the curves are referred to as the optimum family of curves. Since a certain amount of flight time will be required by the missile to reach the target, and the earth is rotating, the target is actually moving in space as the missile travels toward it. Because of this a missile traveling a highaltitude trajectory must seek to hit the target at a diiferent point in space than a missile traveling a low-altrtude trajectory. Therefore although all of the curves in the optimum family intercept the same point on the earth (is. the target) they would not all intersect at the same point in space.

After the missile has been launched, and accelerated to a position and velocity approaching those required to follow one of the optimum family of curves, with the radar tracking the missile, the computer continuously calculates the missiles trajectory. At the same time, the computer continuously tests to see if the missiles trajectory (i.e., position and velocity) will satisfy the mathematical equation of any of the optimum family of curves, and

if not, calculates the error or disparity between it and the nearest'curve of the family. This disparity is then resolved into control signals by a command link and transmitted, via command links communications system, to the missile. The command links communications can be effectively established by providing for side band modulation of the transmitted radar signals and equipping the missile with a signal discriminator and aside band recciver. Afiight control system 23 is provided in the missile, as will be described later, to control its velocity and this fii ht control system converts the control signals communicated to the missile by the command links into regulation'ol the missiles velocity so "as to reduce the calculated disparity.

in this manner, one of' the optimum family of curves is selected as curve 12, and the missile is brought to the initial conditions required to cause a free mass to follow that curve. At this instant the computer indicates a solution and a suitable timed command signal is sent to the missile via the command links communication'to activate its self-contained inertial guidance system 24.

If, in seeking to bring the missile to the initial conditions (i.e., position and velocity) required by one of the optimum family of curves, those conditions required by another curveare attained first, the computer will indicate a solution with that curve instead, and activate the missiles inertia guidance System24- at that point. Thus,'the computer, will put the 'rnissil on the first curve in the optimum family whose equation is satisfied by the missiies position and velocity. This simplifies the problem of attaining the initial conditions byproviding a plurality of possible solutions instead of only one. Also, it should be understood that it is not necessary for the position and velocity of the missile to meet the requirements of a curve exactly, but only that the calculated disparity be rcduced until it i within tolerable limits; Thisalso sir-nplifies the problem of attaining the initial conditions.

Ground control systems of this type are known in the art as indicated on page3l of Burroughs Corporation publication entitled Weapons Systems Management l paper at theN ALTO. Seminar (referenced supra) isalso pertinent, as is a paper entitled Digital Techniques for Missile Guidance Systems by Sidney Darlington pre.

"sented at'the' same Seminar.

With regard to guiding the missile onto a free mass -in FIGURE 3.

trajectory curve, it should also be kept in mind that since the computer 21 is located on the ground, it can be made as large and as accurate as necessary in order to obtain the desired results. The computer in my system, being ground-based, is not subject to the design limitations of the missile-borne computers used in some prior known guidance systems.

When the missile has been guided onto a suitable curve 12, and the inertial guidance system actuated, the function of the ground control system 17 is still not necessarily completed. The ground control system 17 may then become a monitor and continue to track the missile and compare its trajectory with the equation of a chosen curve 12 to be sure that its solution was correct and that the missiles self-contained inertial guidance system is operating properly. If th inertial guidance system is not functioning properly, or if it is desirable to further reduce the calculated disparity and more accurately position the missile on the curve, the inertial guidance system. 24 can be overridden, as will be explained later, and the missile controlled again from the ground.

if, after a specified period of monitoring during which the necessary adjustments can be made, as indicated above, the computer finds that the missile is properly following the selected curve 12, a second command signal is sent to the missile which arms its explosive charge and prevents all further external control to avoid enemy jamming. if, on the other hand, the computer determines after continuous trials and adjustments, that the inertial guidance system is not operating properly, or for some other reason the missile is malfunctioning, the computer will issue another command signal to detonate destruction charge and destroy the missile in air, The missile is therefore continuously monitored until it is determined to be operating properly.

lt will also be appreciated that, if desirable, the missile can be equipped with a self-contained flight programmer of inertial variety of controlthe missile during its initial stage and the ground control system utilized only as a monitor. Such a flight programmer would not be required to meet extreme accuracies and would only function for a short period, and therefore could be constructed relatively inexpensively.

From this description of my ground control system, it will be seen that I have provided a very effective method for establishing the proper initial conditions required for my missile to follow a chosen free mass trajectory curve in accordance with the the above-mentioned principle of operation of my invention. It should be understood, however, that other methods of initial stage control, such as guidance from aircraft or ships, might also be used, the essential feature being the establishment or" the proper initial conditions.

lldisrile Components The missile itself and its major components are shown At its nose, the missile has a warhead compartment which carries the explosive 25, and, just behind this, a radar compartment which contains the radar repeater 22 for receiving and retransmitting signals from the ground transmitter 229.

Next, is a compartment containing radio equipment 3d used to establish the command links communications, such as the side band receiver and signal discriminator.

Behind the radio equipment 26 is a control compartment 27 in which a portion of the flight control means, previously referred to, is located.

At the center of the missile is a guidance system compartment 28 which contains a portion of the inertial guidance system 24 that maintains the missile on the curve 12 during the latter stages of its flight. A body 29 having a fixed mass is part of the inertial guidance system 2 3- and is contained in the guidance system compartment 23.

Aft of the guidance compartment is an auxiliary power compartment which contains an auxiliary power source 30 for supplying power to the missiles control apparatus, and in the tail section, a propulsion compartment is provided which contains the missiles propulsion means 31.

The propulsion means 31 may be any suitable rocket or jet propulsion engine. In fact, it is contemplated that propulsion might be provided in three stages: a ramjet initial and final stage, and a middle stage in the rarer atmosphere during which the oxidizer is provided partly or entirely from tanks within the missile.

To provide means for steering the missile in flight, aerodynamic control surfaces 32, 33 and 3d are mounted on the tail 35. T he control surfaces each have deflectors 36, 3'7 and 33 which extend inwardly over the exhaust pipe of the propulsion unit 31 and assist the steering of the missile by deflecting the jet stream. The deflectors are particularly necessary in rarefied atmosphere where the effect of the aerodynamic control surfaces is greatly reduced. Deflectors of this type were used for steering the German V-2 rocket and are described in United States Pat nt 2,644,296, dated luly 7, 1953, and entitled Laminated let Vane.

I also contemplate, as part of the missile guidance system, an aerodynamic flight stabilization control 3d, which controls the roll, yaw and pitch of the missile, either independently or in cooperation with the inertial guidance system 24. The flight stabilization control 39 has two gyros for controlling roll, pitch and yaw. Flight stabilization systems of this type are known in the art as indicated in the book Naval Ordnance and Gunnery, vol. 2, published by the Department of Ordnance and Gunnery, US. Naval Academy (1955), particularly at page 379 et seq.; the Locke reference (supra), particularly at page 46, FIGURES 215; and in the Air Force book Guided Missiles (supra), particularly at page 362 et seq.

in FIGURE 4, I show a block diagram of the components of my missile. During the initial stage of flight when the missile is guided by my ground control system 17, the signals from the radar transmitter 20 are received by antenna 41, carried via a signal discriminator 42 to radar repeater 22, and then re-transmitted. The side band signals of the command link are also received by antenna 41 and are carried via signal discriminator 42 to side band receiver 43 where they are amplified and passed to control circuits id. The control circuits dd convert the command link signals into power signals which motivate servo actuating means 45 to regulate the missiles propulsion means 31 and control surfaces 32, 33 and 34. The control circuits 44, actuating means 45, and the flight controls (propulsion means and control surfaces) comprise the flight control system 23, previously mentioned.

When the command link signals to activate the inertial guidance system, or arm or destroy the missile, the signals are passed by the side band receiver 43 to release device 46 of the guidance system 24, arming device 47, or destruction device 48, respectively.

When guidance has been turned over to the inertial guidance system 24, the missile is controlled by a followup system 50 which responds to signals from a guidance detector 49. These components will be discussed in detail later.

Power for this missile-borne equipment is supplied by an auxiliary power unit 51.

Inertial Guidance System In FIGURE 5, I show a detailed View of the guidance detector 49 of my inertial guidance system 24. The detector is shown schematically and consists principally of aoraseo tracts to release body 29 at a reference position within the housing 6! in which is mounted an axially movable arm 61 with a cup-shaped head 62 on the outer end thereof. The housing contains an electric coil which forms an electromagnet with the arm 61 such that when the coil is energized the arm moves to an extended position (see FIGURE and when tie-energized, thearm retracts. The cup-shaped head 62 has a seat designed to hold the body 749, and thearrn 61 is of such a length that when extended the head holds the body in its reference position in the chamber 62.

The lock 53b secures the body 29 in the head 62 of sup port 53a during the initial stage of the missiles travel. It consists of a housing 65 which has a pin 66 mounted therein with one end extending therefrom. The housing 65 contains an electric coil which forms an electromagnet with the pin 66 and when energized causes the pin 66 to move outwardly to an extended position with respect to the housing. When the coil is de-energized the pin 66 returns to a retracted position.

The lock 53b is positioned so that when pin 66 is in its extended position it bypasses the body 29 on the side. opposite to the head 6?; in such a manner that the body 29' can be contained between the pin es and the headtiZ when the coils of both the support 53a and lock 53b are energized, and held at the reference position. Since the greatest acceleration forces to which the body 29 is subjected while held by the supporting device will take place when the missile is in a nearly vertical position, they will be concentrated primarily on support 53a and directed substantially axially with respect to the arm til. With proper construction the support could easily withstand such forces.

Release. of the body 29 when thus contained, is accomplished by first derenergizing the coil of the lock 53b to retract pin es, and then shortly thereafter, de-en rgizing the coil in support 53a tozretract the arm 61 and head 62.

With this arrangement the seat in head 62 prevents the body 2 9 from following the pin 66 as it is retracted and then frees the body without imparting any new forces to it when arm 61 retracts. From this description it will be seen that when released body 29 is freely disposed in the chamber, and, so long as it does not' contact the sides of the chamber, it acts like a freely falling body.

The body 29 is therefore free to act just as my previously mentioned theoretical mass-body. outside influences, except. gravity andits previously imparted velocity, it proceeds, in accordance with the laws of celestial mechanics, along the free mass trajectory curve 12 toward the target.

Isolated from all In order to cause the missile to follow the same trajectory as the body 29, the missile is slaved to the body by means of detection devices which sense any displacement of the walls of chamber 52 With respect to body 29 and the follow-up system 50, which, via the previously mentioned flight control means, corrects the velocity of the missile so as to minimize this displacement and keep the missile at the reference position with respect to the body. i

The detection or pick-off devices 55, 56. and 5'7 are hi ghly sensitive to changes in electrical capacitance, and the body 29 is made of metal so that its distance from the devices affects their charge. The detection devices are located one on each of the three coordinate axes of the reference position, which is located at the center of. the missile, as shown, and equidistant therefrom, andtherefore give a continuous indication of the position of body 2 19. When thebody is released by the supporting device 53, the. chamber 52 will be in. the-referenceposition, Position l as shown in FIGURE 5.

The follow-up system 5% has an amplifier and analyzer which amplify and analyze the signals from the detection devices and produce resultant signals to the flight control system 23 which change the missiles velocity so as to keep chamber 52 in reference position 1 with respect to body 2?. in the reference position, theoutput signals of the detection devices 55, 56-and 57 are balanced and therefore no resultant signal is produced to vary the missiles velocity.

If no forces other than gravity were acting on the missile, it would remain in reference position 1 with respect to the body 29 and travel along curve 12 to target ll without further guidance. This is not the case, however, because the missile is subjected to the atmosphere,and therefore to atmospheric forces which must be compensated for if the missile is to stay on curve 12.

One such force is atmospheric drag. Since theinertial.

guidance system doesnt take over until after the initial conditions for curve 12 are satisfied, in the case of long range missiles this time lag may be such that by the time the guidance system takes over, the atmospheric drag has very'little effect because of the rareiicstion of the atmos phere at high altitudes. This will not necessarily be true for medium range missiles, however, and, in any case, all types of missiles must re-enter the denser atmosphere to reach a target on the earths surface so that, at least during this final stage, atmospheric drag will be a significant force to be overcome.

Atmospheric drag when effective will continuously tend to slow the missiles speed and therefore tend to move it backward with respect to body 29 to position 2 (see FIG- URE 6). In this position, the detection device 56 will be closer to body 29 than in position 1 and will indicate this change by an appropriate signal. Detection devices 55 and 57 will each be slightly farther away from the body 219 than in position 1, but still equidistant therefrom and will indicate their situation by appropriate signals.

The signals of all three of the detectors are analyzed by an analyzer circuit in the follow-up system 58 and resulting correction signals are produced which change the missiles guidance controls so as to return it to position 1 with respect tobody 29. In this case, where the missile is at position 2, the analyzer will call for more thrust to overcome the drag.

It will be appreciated, of course, that if the conditions of flight will be such that the atmospheric drag will at some points be so reduced that negative thrust might be necessary, reverse propulsion means can be provided'for this purpose.

Another atmospheric force which must be corrected for is wind. During the missiles course of travel, it will be subjected to unpredictable winds which will tend to drive it oif its path and will therefore move the missile with respect to body 29. if, for example, a Wind strikes the missile from an angle of degrees relative to its heading and directed 30 degrees downwardly from horii zontal, it will tend to move the missile to position 3 with respect to the body 29 (see FIGURE 7). In this position,

the body 2& is farther from all three detectors than when in position 1, and signals from the detection devices will indicate this changerof proximity. The analyzer circuit of the follow-up system will analyze the signals from the detection devices and produce resultant signals to the control circuits 44 calling for a reduction of thrust and an appropriate change in heading to compensate for the wind.

Actually, of course, the control exercised by inertial guidance device 24 willbe substantially instantaneous and continuous, so that, in reality, the missile, instead of being driven off curve 12 by extraneous influences and then guided back on, is kept on course by continuously and substantially instantaneously compensating for extraneous influences which tend to displace it.

The detection devices 55, 56 and 57 are described as electro-capacitance type proximity pick-offs, but it will be appreciated that any other type of detector might be used which provides an instantaneous indication of the proximity" of body 29 without applying any appreciable coercive force to it. The body 29 could be made luminescent, for instance, and photo-electric cells used as detection devices, or the body 29 could be charged and electrostatic detection devices used.

As previously mentioned, my inertial guidance system 24 is capable of being overriden by the ground control system 17 during the monitoring stage of its flight. To permit this, the supporting device 53 in my guidance detector 49 is capable of momentarily recapturing the body 29 and holding it fixed during the periods of overriding.

The mass of body 29 with respect to the missile may differ depending on the inertial characteristics desired. However, the basic principles of operation apply irrespective of this relationship.

It will also be appreciated that in order to prevent sudden forces from driving the missile far enough off course to cause the walls of chamber 52 to engage body 2%, the response time of the missile control system must be relatively small. This requirement is not intolerable, however, because the mass of the missile is large enough to greatly dampen the effect of sudden forces on it, and chamber 52 can be made large enough with respect to the body 29 to permit considerable movement therebetween. Also, if the missile should encounter a force sufficiently strong and sudden to cause chamber 52 to engage the body 29, unless the force is prolonged, the error introduced, not being cumulative, would still normally be quite tolerable. That is, the errors tend to balance out over a trajectory on a normal statistical basis.

To assure that the detected movement of the missile with respect to body 2h is properly corrected by the missile controls, it is also necessary that the response time of the detection devices and follow-up system be low with respect to that of the flight stabilization control 39.

Since a major portion of target error is due to unpredictable variations in the atmospheric conditions encountered by the missile, my inertial guidance system significantly increases the missiles accuracy by causing it to automatically correct for these conditions. The only remaining sources of error are the residual error in the 1 computer and other equipment, inaccurate knowledge of the target or launch point coordinates, anomalies in the earths gravitational field along the missiles path, and unknown variations in refractive index of the atmosphere above the launch point which affect the accuracy of the radar.

in general, the errors introduced by these sources are inherently small or are controllable and can be kept quite small. Considering the residual error in the ground control 17, for example, radar experts claimed several years ago that velocity components could be measured to 0.2 foot per second. If this were the only appreciable error, a missile utilizing my guidance system could strike a target at a range of 400 miles with a range error of only 141 feet and a lateral error of only 36 feet.

Understanding the operation of my inertial guidance system, it will be appreciated that my guidance system is SllfilClCHllY simple, when compared to prior systems, to permit its production at a considerably lower cost, particularly since the simplicity of my system would allow the use of mass production techniques and relatively unskilled labor to a much greater extent. Furthermore, the simplicity of my system increases its inherent reliability since the amount of missile-borne precision equipment required is substantially reduced. Because of these important features, my system makes the guidance of missiles with conventional warheads, as well as those with atomic warheads, economically feasible;

Though the system which I have described controls a ground-to-ground missile, it will be appreciated that my invention could also be advantageously used for an air-toground missile and in other guidance applications.

While the forms of my invention, herein disclosed, are fully capable of attaining the objects and providing the advantages heretofore described, it should be underit) stood that I do not mean to limit myself to the particular details disclosed except as provided in the appended claims.

I claim:

l. The method of guiding a missile having flight control means capable of regulating its velocity which comprises: disposing a body having mass in said missile; imparting to said body and missile certain initial conditions of velocity and location corresponding to requirements of a position on one of a family of curves which represent ballistic trajectories of a free mass in a vacuum; isolating said body from the influence of all forces except gravity; and slaving said missile to said body by regulation of said flight control means so as to minimize any relative movement therebetween.

2. The method of guiding a missile, which comprises the steps of: placing a body of finite mass inside a protected chamber provided Within the missile and in rigidly supported relationship therewith; applying propelling and guiding forces to the missile for a finite period of time until certain desired conditions of velocity and location are achieved; removing the rigid support of said body so that said body is free to move within said chamber under the influence of gravity; detecting relative motion between said chamber and said body; and applying propelling and guiding forces to said missile so as to minimize the relative motion between said body and said chamber.

3. In a guidance system for a vehicle which has flight control means for regulating the velocity of said vehicle, and initial stage control means for controlling said vehicle through said flight control means to establish certain initial conditions of velocity and location, an inertial guidance system for controlling said vehicle after establishment of said initial conditions comprising: a reference frame associated with said vehicle; a mass; releasable means for holding said mass fixed in said reference frame in a reference position, said mass being released upon establishment of said conditions to follow a predetermined path when acted upon by known forces; isolation means within said missile for isolating said mass from all but said known forces; and guidance control means in said vehicle con nected with said flight control means for measuring displacement of said mass from said reference position with respect to said reference frame and regulate said flight control means to reduce said displacement whereby said vehicle is caused to follow said mass.

4. Inertial guidance means fora vehicle having flight control means for regulating its velocity comprising: a chamber in said vehicle; a body having mass positioned in said chamber; releasable support means for holding said body fixed with respect to the walls of said chamber in a reference position, said support means being actuat able to release said body and permit said body to move freely in said chamber; detection means in said vehicle for sensing relative displacement between said body and said reference position; and follow-up means intercoupled between said detection means and said flight control means for converting the output of said detection means into regulation of said flight control means to minimize said relative displacement between said body and said reference position.

5. In a guidance system for a missile which has flight control means in said missile for regulating the velocity of said missile, and initial stage control means on the ground which control said flight control means during the initial stage of said missiles flight until certain initial conditions of velocity and location are established, an inertial guid ance system in said missile for controlling said flight control means after establishment of said initial conditions comprising: an evacuated chamber; a body having mass positioned in said chamber; a releasable supporting device in said chamber for fixedly holding said body therein until said initial conditions are established, said supporting device being actuatable to free said body in a reference position in said chamber spaced from the walls thereof;

proximity detection devices mounted in said missile and disposed about said reference position, said detection devices having an output indicative of their proximity to said body; and follow-up means intercoupled with said detection devices and said flight control means for converting the output of said detection devices into regulation of said flight control means so as to maintain said body in said reference position.

6. A guidance system for a vehicle having flight con trol means for controlling the velocity thereof, comprising: means for regulating said flight control means t bring said vehicle to certain initial conditions of velocity and location; a body having mass within said vehicle; means for fixedly supporting said body in said vehicle until said initial conditions are established; means for isolating said body from all forces except gravity after said initial conditions have been established; and guid ance control means in said vehicle connected to said flight control means and responsive to forces acting on said vehicle other than gravity for regulating the velocity of said vehicle to cause it to follow said body.

7. A guidance system for a missile having flight c ntrol means for controlling the velocity thereof, com prising: initial stage control means for controlling said flight control means to bring said missile to certain initial conditions of velocity and location required to follow one of a particular family of paths which represent gravitational trajectories of a free mass in a vacuum that intercepts a selected target; an evacuated chamber in said missile; a body having mass positioned in said chamber; means for fixedly supporting said body prior to establishment of said initial conditions; means for releasing said body in a reference position spaced from the walls of said chamber when said initial conditions have been established; detection means for detecting movement of said missile with respect to said mass after the release thereof; and follow-up means intercoupling said detection means and said flight control means, and responsive to the output of said detection means to regulate said flight control means and guide said missile so.

as to maintain said body in said reference position.

8. A guidance system for a missile comprising: flight control means in said missile including thrust producing means; initial stage control means including groundbased radar and computer means for calculating the disparity between the trajectory of said missile and the nearest one of a selected family of curves which represent gravitational trajectories of a free mass in a vacuum that intercept a selected target, and command means for regulating said flight control means to minimize said disparity; an evacuated chamber Within said missile; a body having mass positioned in said chamber; means for fixedly supporting said body in said chamber until said disparity is reduced to a predetermined amount;

means for thereafter releasing said body in said chamber spaced from the walls thereof; and guidance control means in said missile connected to said flight control means and responsive to relative movement between said body and said chamber walls to cause said missile to follow said body after release thereof.

9. A guidance system for a missile comprising: flight control means in said missile including thrust producing means and control surfaces for propelling and steering said missile; radar tracking means including a groundbased transmitter and receiver for tracking said missile during the initial stage in its flight; a ground-based computer in communication with said radar tracking means for calculating the disparity between the trajectory of said missile and the nearest one of a selected fam ly of curves which represent gravitational trajectories of a free mass in a vacuum that intercept a selected target, and indicating a solution when said disparity is reduced to a predetermined amount; a command link coupled to said computer and communicating between said computer and said missile for regulating said tight control means so as to reduce said disparity and for transmitting a re- .means and flight control means which convert the output of said detection means into regulation of said flightcontrol means to change the velocity of said missile so as to minimize said relative movement.

10. A guidance system for a missile comprising flight control means in said missile for controlling the velocity thereof; means for regulating said flight control means to bring said missile to certain initial conditions of velocity and location required to follow one of a particular family of paths which represent gravitational traiectories of a free mass in a vacuum that intercept a selected target; an avacuated chamber in said missile; a metal body having mass positioned in said chamber; releasable holding means mounted in said chamber for fixedly holding said body with respect to said chamber until said initial conditions are met and thereafter releasing said body in a reference position spaced from the walls of said chamber; three detection devices each having an electric capacitance which varies with the proximity of said body and being disposed in said chamber equally spaced from said reference position with one device aligned with a longitudinal axis through said reference position which is parallel to the longitudinal axis of said missile, a second device aligned with a vertical axis through said reference position which is parallel to the vertical transverse axis of said missile, and the third device aligned with a horizontal axis through said reference position which is parallel to the horizontal transverse axis of said missile; and followup means intercoupled with said detection devices and flight control means for converting changes in the capacitance of said detection devices into signals for regulating said flight control means to change the velocity of said missile so as to maintain said body in said reference position.

11. A guidance system for a missile comprising: flight control means in said missile for controlling the velocity thereof; means for regulating said flight control means to bring said missile to certain initial conditions of velocity and location required to follow one of a particular family.

of paths which represent gravitational trajectories of a free mass in a vacuum that intercept a selected target; an evacuated chamber in said missile; an electrostatically charged body having mass positioned in said chamber; releasable holding means mounted in said chamber for fixedly holding said body with respect to said chamber until said initial conditions are met and thereafter releasing said body in a reference position spaced from the walls of said chamber; a plurality of detection devices which are sensitive to an electrostatic charge for producing an output indicative of their distance from said electrostatically charged body, said detectors being mounted in said missile at different locations about said reference position but equally spaced therefrom; and follow-up means intercoupled with said detection devices and said flight control means for converting the output of said detection devices into signals for regulating said flight control means to vary the velocity of said missile so as to maintain said body in said reference position.

12. A guidance system for a missile comprising: flight control means in said missile for controlling the velocity thereof; means for regulating said flight control means to bring said. missile to certain initial conditions of velocity and location required to follow one of a particular family of paths which represent gravitational trajectories of a 13 free mass in a vacuum that intercept a selected target; an evacuated chamber in said missile; a light-emitting body having mass positioned in said chamber; releasable ho1ding means mounted in said chamber for fixedly holding said body with respect to said chamber until said initial conditions are met and thereafter releasing said body in a reference position spaced from the Walls of said chamher; three photoelectric cells mounted in said missile equally spaced from said reference position and exposed to the light from said body with one of said cells aligned with a longitudinal axis through said reference position which is parallel to the longitudinal axis of said missile, a second of said cells aligned with a vertical axis through said reference position Which is parallel to the vertical transverse axis of said missile, and the third of said cells aligned with a horizontal axis through said reference position which is parallel to the horizontal transverse axis of said missile, said photoelectric cells each having an electrical output indicative of their distance from said lightemitting body; and follow-up means intercoupled with said cells and said flight control means for converting the output of said cells into signals for regulating said flight control means to vary the velocity of said missile so as to maintain said body in said reference position.

References Cited in the file of this patent UNITED STATES PATENTS 2,603,433 Nosker July 15, 1952 2,603,434 Merrill July 15, 1952 2,916,279 Stanton Dec. 8, 1959 2,932,467 Scorgie Apr. 12, 1960

Claims (1)

1. 3. IN A GUIDANCE SYSTEM FOR A VEHICLE WHICH HAS FLIGHT CONTROL MEANS FOR REGULATING THE VELOCITY OF SAID VEHICLE, AND INITIAL STAGE CONTROL MEANS FOR CONTROLLING SAID VEHICLE THROUGH SAID FLIGHT CONTROL MEANS TO ESTABLISH CERTAIN INITIAL CONDITIONS OF VELOCITY AND LOCATION, AN INERTIAL GUIDANCE SYSTEM FOR CONTROLLING SAID VEHICLE AFTER ESTABLISHMENT OF SAID INITIAL CONDITIONS COMPRISING: A REFERENCE FRAME ASSOCIATED WITH SAID VEHICLE; A MASS; RELEASABLE MEANS FOR HOLDING SAID MASS FIXED IN SAID REFERENCE FRAME IN A REFERENCE POSITION, SAID MASS BEING RELEASED UPON ESTABLISHMENT OF SAID CONDITIONS TO FOLLOW A PREDETERMINED PATH WHEN ACTED UPON BY KNOWN FORCES; ISOLATION MEANS WITHIN SAID MISSILE FOR ISOLATING SAID MASS FROM ALL BUT SAID KNOWN FORCES; AND GUIDANCE CONTROL MEANS IN SAID VEHICLE CONNECTED WITH SAID FLIGHT CONTROL MEANS FOR MEASURING DISPLACEMENT OF SAID MASS FROM SAID REFERENCE POSITION WITH RESPECT TO SAID REFERENCE FRAME AND REGULATE SAID FLIGHT CONTROL MEANS TO REDUCE SAID DISPLACEMENT WHEREBY SAID VEHICLE IS CAUSED TO FOLLOW SAID MASS.

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