US3456512A - Inertial device utilizing gyroscopes - Google Patents

Description

July 22, 1969 R. L. SCHMIDT IIIEIFUIIAL DEVICE UTILIZING GYROSGOPES 3 Sheets-Sheet 1 Filed Sept. 13, 1967 mv mm INVENTOR Robert L. Schmidt I M K BY gofim im y 969 R. L. SCHMIDT INERTIAL DEVICE UTILIZING GYROSCOPES 3 Sheets-Sheet 3 Filed Sept. 13, 1967 United States Patent U.S. Cl. 74-5.34 6 Claims ABSTRACT OF THE DISCLOSURE An apparatus to monitor the position of an accelerating missile relative to the launching point with respect to a desired straight line trajectory, and to give an arming signal if the missile is within the desired trajectory of a safing or destruct signal should the missile deviate from the desired trajectory. A pair of displacement monitoring gyroscopes having movable mass elements provide an indication of deviations in the desired trajectory about the pitch and yaw axes of the missile in reference to a third solid mass gyroscope that cancels the roll angle from precession experienced by the displacement gyroscopes.

Background of the invention This invention relates generally to inertial guidance systems and more particularly to a mechanical inertial system to provide missile position intelligence with respect to a desired straight line trajectory.

Missiles launched for a straight line trajectory toward a target may encounter a deviation from the desired trajectory due to a failure within the missile or some disturbing external force acting on the missile. A deviation from the desired trajectory will cause the missile to miss the selected target, such that the missile may strike an undesired target. It has been found desirable to provide devices which monitor the missile trajectory while in flight and to arm the missile only if on target course or to render the missile ineffective if off target course.

Fixed longitudinally accelerometers have been employed to provide environmental safety by arming the missile only upon sensing a predetermined acceleration, but these devices will function even though the missile is on an unwanted trajectory.

More complex, electro-mechanical guidance systems have been used occasionally to provide position intelligence, but these devices require complex electronic networks to integrate the displacement in three dimensional space. Such systems are capable of monitoring and controlling the trajectory, but they are more precise and more expensive than is warranted for the usual arming or safing functions.

Summary of the invention Accordingly, an object of the present invention is to provide a new and improved displacement monitoring gyroscope.

Another object of this invention is to provide a gyroscope which can monitor and indicate a deviation of an aerial vehicle from a straight line trajectory.

A further object of the instant invention is to provide a gyroscope which monitors missile trajectory and indicates missile position with regard to a desired trajectory.

Astill further object is to provide a mechanical inertial system to provide missile position intelligence with respect to a desired straight line trajectory.

Briefly, in accordance with one embodiment of this invention the foregoing and other objects are attained by providing an inertial system utilizing displacement monitoring gyroscopes which produces mechanical output signals to indicate deviations from a desired trajectory of a "ice missile in flight about the pitch and yaw axes, and arms the missile if on the acceptable trajectory or renders the missile safe if it has deviated from the acceptable trajectory.

Many other advantages and objects of the present invention will become readily apparent as the following detailed description of the invention unfolds when taken in conjunction with the accompanying drawings.

Brief description of the drawing FIG. 1 is a schemtic illustration of a missile in launch position and the acceptable cone of trajectory;

FIG. 2 is a perspective illustration of the inertial safing and arming system of the present invention;

FIG. 3 is a fragmentary plan view showing the flyball linkage arrangement of the displacement gyroscope;

FIG. 4 is a perspective view of the Y-displacement gyroscope during missile nose down; and,

FIG. 5 is a schematic wiring diagram of the central circuitry.

Description of the preferred embodiment Referring now to the drawings wherein like reference characters designate identical or corresponding parts throughout the several figures and more particularly to FIG. 1 whereon a missile indicated schematically at 1 is mounted on a launching mechanism 2. The desired straight line trajectory for the missile is indicated by the line 3, while the acceptable limits of trajectory are indicated by the cone 4. The apex 5 of the acceptable cone of trajectory is located on the longitudinal axis of the missile where an arming or safing device 6 in accordance with the instant invention is located. The arming or safing device 6 will be described in detail hereinafter but for the moment it may be assumed that it operates to sense a deviation in trajectory from a normal desired straight line 3 and provides an output signal arming the missile if within the cone 4 or an output safing signal disarming or destroying the missile if outside the cone 4. The particular mechanisms for arming and/ or safing the missile are not part of the instant invention and any conventional mechanism may be employed with the invention.

As shown in FIG. 2, the arming and safing device 6 consists of an intertial or reference platform having a plurality of gyroscopes 7, 8 and 9 for respectively sensing yaw or Y displacement, roll and pitch or X displacement of the missile and an accelerometer 11 the purpose of which will be explained more fully hereinafter. Four trunnions 12, 13, 14 and 15 are secured to the missile such that the pivotal axis of each trunnion corresponds to the longitudinal axis Z of the missile. The Y displacement gyroscope 7 is pivotally mounted on the trunnions 12 and 13 by pins 16 and 17'. The pin 16 is fixedly secured to the outer gimbal 18 of the gyroscope 7, and is pivotally mounted in a bearing (not shown) in trunnion 12. Pin 17 is fixedly secured to the outer gimbal 18 opposite to pin 16 and in alignment therewith. Similarly the outer gimbal 24 of the X displacement gyroscope 9 is pivotaly mounted by pins in trunnions 14 and 15.

Pin 17 extends through a bearing (not shown) in trunnion 13 and includes an indicator or switch finger 19 integral therewith. The pin '17 further extends through a stationary disc 21 having a conductive strip 45 centrally disposed thereon and the outer gimbal 22 of the roll gyroscope 8 such that the roll gyroscope is pivotally mounted thereon and on a pin 25. Elements 19 and 21 constitute a rotary electrical switch. The disc 21 is fixedly secured to the outer gimbal 22 as is a rotary switch disc 23 on the opposite side thereof, The central roll gyroscope 8 includes a balanced rotating cylindrically shaped solid mass 26 mounted by means of pins or bearings 27 for rotation within an inner gimbal 28. The inner gimbal 28 of the roll gyroscope 8 is mounted for rotation with respect to outer gimbal 22 by means of pins or bearings 29. Since the roll gyroscope 8 is a solid mass stable gyroscope, the axis of rotation of mass 26 will always be oriented parallel to the Y axis of the missile regardless of its orientation. Furthermore, the outer gimbal 22 of roll gyroscope 8 will never precess or rotate, and therefore the roll gyroscope 8 prevents a fixed reference for the X displacement gyroscope 9 and the Y displacement gyroscope 7 while the missile experiences roll. Since gyroscopes 7 and 9 are structurally identical and merely oriented ninety degress out of alignment about the missile Z axis, only the structure and operation of the Y displacement gyroscope 7 will be described in detail.

An inner gimbal 31 is mounted for rotation within outer gimbal 18 by means of pins or bearings 32. A cylindrically shaped mass 33 is rotatably mounted in gimbal 31 by means of bearings 34 and rotates at an angular velocity to. If the mass 33 were solid and balanced, the Y displacement gyroscope 7 would be identical to the roll gyroscope 8. However, mass 33 is not solid, but has a hole 35 radially extending completely therethrough which intersects the axis of rotation of mass 33. Perpendicular to hole 35 is a second hole 36 the axis 'of which also intersects the axis of rotation of mass 33. Hole 36 terminates at a wall 37, at which point a pin guide hole 38 extends through the center of the mass 33. A similar structure exists on the other side of the center of mass 33.

Referring specifically to FIGS, 2 and 3, it can be seen that a pair of sensing masses such as balls 39 and 41 are pivotally mounted within hole 35 by means of links 42 and 43. Links 42 and 43 are pivotally anchored to a pin 44 disposed in pin guide holes 38. Hole 35 is slightly larger in diameter than the diameter of balls 39 and 41 and of sufficient length to permit the balls to move above or below the center of mass of the mass 33.

In order to better understand the overall operation of the fiyball gyroscope, an analysis will be given of the gyroscope under normal and abnormal trajectory conditions and a comparison made of the gyroscope reaction thereof under both conditions. Regardless of the particular trajectory the missile is following, whether normal (within the cone defined by line 4 shown in FIG. 1) or abnormal (outside the cone shown in FIG. 1), the axis of rotation of the roll gyroscope 8 will always correspond to the Y-axis as illustrated in FIG. 2. Thus the roll gyroscope acts as an inertial reference for both the X and Y displacement gyroscopes.

Referring to FIG. 2 it can be seen that the rotating mass 33 in the Y displacement gyroscope 7 produces an angular spin velocity :0. The angular momentum vector H for a rotating gyroscope is equal to the product of the angular spin velocity to and the moment of inertia I of the rotating mass. (H=Iw). The direction of the angular momentum vector H is found by the right-hand rule. Thus, wrapping the fingers of the right hand around the rotating mass 33 in the direction of rotation, the thumb will point in the direction of the angular momentum vector H. As can be seen in FIG. 2 a counter-clockwise rotation to produces an angular momentum vector H, which points straight up along the axis of rotation of the Y displacement gyroscope.

When the missile is travelling in a normal straight line trajectory, there will be two forces exerted on balls 39 and 41. As shown in FIG. 2, there is a centrifugal force F exerted on each of balls 39 and 41 due to the rotation of mass 33 and balls 39 and 41, which force acts on the balls to throw them outwardly and centers the balls in the slots 35. There is also a setback force F exerted on each ball due to the acceleration of the missile. In a normal straight line trajectory along the Z-axis, the

setback force F will be exerted on the balls parallel to the Z-axis and in a direction opposite to the direction of missile travel. Since, all the forces exerted on the Y displacement gyroscope 7 when the missile is in a normal straight line trajectory lie on the axis of the gyroscope, there are no resulting moments acting on the gyroscope and thus there will be no gyroscopic precession. Since no gyroscopic precession results from a normal trajectory, the indicator or switch finger 19 will be centered on the rotary switch disc 21.

Should the missile either nose down or up away from a normal trajectory, there will be precession of gyroscope 7 and the indicator or switch finger 19 will move offcenter relative to the rotary switch disc 21. The above result can be seen from an analysis of the forces acting on the gyroscope as the missile noses down. The axis of rotation of the outer gimbal 18 of the Y displacement gyroscope will always lie in the same plane as the missile axis, since it is secured to the missile by trunnions 12 and 13. The inside gimbal 31 will always lie in the XY plane in accordance with standard gyroscope principles. As can be seen in FIG. 4, as the missile noses down the outer gimbal 18 will assume an angle a with the Z axis. The angular momentum vector H lies along the Y axis and points straight up in accordance with the right-hand rule. The setback force F2 acting on the balls will lie in a plane parallel to the plane of the missile and outer gimbal 18, and will act in a direction opposite to the acceleration of the missile. Breaking the setback force F2 down into its vector components, it can be seen that there results two forces F3 and F4 acting on each ball. F4 is a force acting in a direction straight up along the Y axis which forces the balls 39 and 41 to the top of their respective slots 35. The force F3 acts in a plane parallel to the Z axis, but above the XZ plane by a distance d. There results a moment M about the X" axis due to the force F3 acting on both balls, equal to twice the product of the setback force component F3 and the distance d(M=2(F3)d). The disturbing moment M acts in a counterclockwise direction about the X axis, as seen in FIG. 4. Applying the righthand rule, a disturbing moment vector M lies along the X axis. The rule of precession for gyroscopes is that the direction of precession of the gyroscope will be the same as the direction the angular momentum vector H will move in if it tries to coincide with the disturbing moment vector M by the shortest possible route. Thus, in accordance with the rule of precession, the direction of precession will be clockwise as the missile noses down and the outer gimbal 18 and the indicator 19 will rotate clockwise. If the missile noses up, the outer gimbal 18 and the indicator or switch 19 will rotate counterclockwise.

The rate of precession Q is equal to the disturbing moment divided by the angular momentum (S2=M/H). As long as the angle of precession does not exceed ((3/2) one half the rotary switch angle 3, the contact 19 stays centered on conductive strip 45 and the switch remains closed. Should the angle of precession exceed 5/2 contact 19 moves off-center and the switch is open. Should some external force such as wind cause the missile to leave the normal cone of trajectory with the missile longitudinal axis remaining parallel to the Z axis, gyroscope precession will still result. Although the setback force F2 will lie in a plane parallel to the XZ plane, the external force (wind, etc.) will cause the balls to move opposite thereto and the resultant moment will cause gyroscopic precession.

It will be evident to one skilled in the art from the foregoing analysis that gyroscope 9 reacts to missile displacement along the X axis, gyroscope 7 reacts to missile displacement along the Y axis and gyroscope 8 functions as a stable reference for gyroscopes 7 and 9 by compensating for missile roll. As long as the missile stays within the acceptable cone of trajectory 4, the switch fingers 19 will remain in contact with the conductive portions 45 of discs 21 and 23 and an arm signal will be provided to the missile arming device. In the event the missile deviates from the acceptable cone of trajectory, the switch fingers 19 will no longer be in contact with the conductive portion 45 and a safe or destruct signal Will be transmitted to the missile.

It is preferable for the safe or arm signal to be given to the missile at a predetermined location along the trajectory. Although the gyroscope system may monitor the missile trajectory from launch to target, it is more desirable to sense a deviation in trajectory and actuate the safe or arming mechanism immediately prior to target deployment.

In order to provide the arming or safing device 6 with the intelligence necessary to locate the missile with respect to distance from the launching device 2 along the Z axis, a conventional accelerometer 11 is provided. For illustrative purposes the accelerometer may be considered to consist of a mass 47 slidingly mounted on guides 48 within a housing 49. By aligning guides 48 parallel to the longitudinal Z axis of the missile, the mass 47 will experience the linear acceleration force of the missile during flight. The mass 47 will move from its rest position against wall 51 of housing 49 toward opposite wall 52, and the rate of movement will be dependent upon a counteracting force exerted by a biasing device such as a coil spring 53 provided between mass 47 and wall 52. A retarding device, such as an escapement 54, is provided to prevent mass 47 from moving too quickly in response to the initial acceleration of the missile in flight or some extraneous external shock when the missile is at rest prior to flight. Three switch contacts 55, 56 and 57 may be mounted on the mass 47, and three corresponding switch contacts 58, 59 and 60 may be positioned on the accelerometer housing 49. Switch contacts 55 and 58 are normally in engagement when the missile is at rest and mass 47 is against wall 51. As shown in the circuit of FIG. 5, upon application of a votlage between terminal T1 and ground G a conventional electroresponsive unlocking device 62 for accelerometer 11 and conventional elect- roresponsive devices 63, 64 and 65 for rotating or spinni ngup the gyroscope masses 26, 33 and 66 are energized. The unlock device may be a solenoid operated latch relay or a magnetic locking mechanism. The spin-up devices for the gyroscopes 7 and 9 are preferably high speed air streams impinging upon the circumference of the masses 33 and 66, but a conventional motor and coil arrangement may be used to rotate all three gyroscope masses. Prior to launching, the spin-up devices bring the gyroscope masses up to full spin. During launch acceleration, the accelerometer mass 47 begins to move rearwardly on guides 48, and electrical continuity between contacts 55 and 58 is interrupted. At some predetermined time during flight, three seconds for example, with the missile experiencing a particular force, such for example as 0.3 g, the mass 47 will have moved a suflicient distance to momentarily close switch contacts 56 and 59, thus operating uncaging mechanisms 67, 68 and 69 for the gyroscopes 7, 8 and 9. The particular uncaging devices are not part of the instant invention and may be any conventional device, for example, a solenoid operated linkage for normally restraining the gyroscope gimbal. From the above discussion it can be seen that prior to missile flight the accelerometer is unlocked and the gyroscope masses are spinning, and during the initial seconds of flight the gyroscopes are uncaged so that they may precess should the missile deviate from the desired trajectory. At some later point in flight the missile will experience a greater g-force, for example 3.5 gs, and the mass 47 will move to aflect engagement of contacts 57 and 60 thus completing the arming or safing circuit between terminals T2 and T3. Should the missile be within the acceptable cone of trajectory, switch fingers 19 will always be in contact with the conductive portions 45 of discs 21 and 23 and the arming or safing circuit will be completed.

Obviously numerous modifications and variations of the present invention are possible in the light of the above teachings.

What is claimed as new and desired to be secured by Letters Patent of the United States is:

1. A gyroscope adapted for use in a missile arming and safing system comprising;

an outer gimbal rotatably mounted on an axis,

an inner gimbal rotatably mounted on said outer gimbal on a second axis perpendicular to said first axis,

a gyroscopic mass rotatably mounted on said inner gimbal on a third axis perpendicular to said first axis and said second axis,

means including at least one pair of sensing masses, each having the same weight and being diametrically opposed equidistant from said third axis, for sensing external forces acting upon said gyroscopic mass for causing gyroscopic precession about said first axis upon sensing forces acting along said third axis,

means pivotally mounting said sensing means on said gyroscopic mass, and

means rotatable with said outer gimbal about said first axis for indicating said gyroscopic precession.

2. An inertial system for providing missile position intelligence with respect to a desired straight line trajectory comprising:

a balanced solid mass gyroscope rotatably mounted on an axis and adapted to provide a stable reference regardless of missile roll,

a displacement gyroscope rotatably mounted on said axis,

means on said displacement gyroscope sensing external forces acting upon said displacement gyroscope for causing gyroscopic precession about said axis,

switch means on said balanced gyroscope having a conductive portion and an insulated portion, and

conductive indicator means on said displacement gyro scope for contacting said switch means,

whereby said indicator means either completes a siganl circuit through said conductive portion or breaks a signal circuit in contact with said insulated portion depending on the degree of gyroscopic precession.

3. The inertial system of claim 2 wherein said displacement gyroscope comprises:

an outer gimbal rotatably mounted on said axis,

an inner gimbal rotatably mounted on said outer gimbal on a second axis perpendicular to said first axis,

a mass rotatably mounted on said inner gimbal on a third axis perpendicular to said first axis and said second axis, and

said sensing means comprises at least one pair of sensing masses each having the same weight and being diametrically opposed equidistant from said third axis.

4. The inertial system of claim 2 including a second displacement gyroscope mounted on said axis on the opposite side of said balanced gyroscope from said first displacement gyroscope and having a spin axis perpendicular to the spin axis of said first displacement gyroscope,

second switch means on said balanced gyroscope having a conductive portion and an insulated portion, and

conductive indicator means on said second displacement gyroscope for contacting said second switch means.

5. The inertial system of claim 4 wherein said displacement gyroscopes each comprise:

an outer gimbal rotatably mounted on said axis,

an inner gimbal rotatably mounted on said outer gimbal on a second axis perpendicular to said first axis,

a mass rotatably mounted on said inner gimbal on a i 7 8 third axis perpendicular to said first axis and said References Cited Second axis, and UNITED STATES PATENTS Said sensing means COIIlPI'ISeS at least 011 6 pair of $6 5- 2 31 455 3 1953 Wing 74 5 34 mg masses each havmg the same welght and be g 301 5 3 57 Wagner 7 diametrically opposed equidistant from said th rd 5 2,821,087 1/1958 Hammon 74--5 .41 axis. 2,963,242 12/ 1960 Mueller 745.34 XR 6. The inertial system of claim 4 including 1g; lliistler a normally open signal circuit electrically connected 1 {m to said switch means and conductive indicator means, 10 3193216 7/1965 Flschel 745'34 XR and 1 t f th 1 t f f FRED C. MATTERN, 111., Primary Examiner an acce erorne er or sensing e acce era lOIl orce o said missile and including means for closing said sig- MANUEL ANTONAKAS Asslstant Exammer nal circuit in response to a preselected acceleration 15 U.S. Cl. X.R.

force. 7 745; 5.6

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