US3245350A - Rocket propelled device for straightline payload transport - Google Patents

Description

April 12, 1966 J. A. KELLY 3,245,350

ROCKET PROPELLED DEVICE FOR STRAIGHTLINE PAYLOAD TRANSPORT Filed April 29, 1963 2 Sheets-Sheet l uIIIIIIIIIIIIIJ FIGURE l 'L Ili...

FIGURE 2 INVENTOR.

ATTY'S April 12, 1966 J. A. KELLY 3,245,350 ROCKET PROPELLED DEVICE FOR STRAIGHTLINE PAYLOAD TRANSPORT Filed April 29, 1963 2 Sheets-Sheet 2 FIGURE 3A FIGURE 38 FIGURE 48 FIGURE 4A INVENTOR. '7 1 A ATTY'S United States Patent 3,245,350 ROCKET PROPELLED DEVICE FOR STRAIGHT. LINE PAYLQAD TRANSPORT Joseph A. Kelly, 2391 Mary St., Riverside, Calif. Filed Apr. 29, 1963, Ser. No. 276,447 1 Claim. (Cl. 102-49) Brief summary of invention Present day infantry weapons which are capable of delivering warheads heavier than hand held rifle fire generally consist of either hand grenades, rifle grenades, mortars, tube launched rockets or recoilless rifles. These conventional infantry weapons suffer deficiencies due to a combination of two or more of the following factors:

( 1) Weapon recoil (2) Inadequate warhead Weight (3) Insuflicient range (4) Excessive size and weight of launcher (5) Trajectory is seriously aflected by atmospheric and gravity forces.

These deficiencies adversely affect the weapon accuracy and dispersion, complicate logistics, limit their application and thereby reduce the weapon usefulness. My invention, as described herein, will provide a warhead delivery capability which will eliminate all or most of the above deficiencies in existing infantry weaponry. In addition, the weapon can be adapted for use as aircraft, land or Water vehicle mounted armament.

Brief description of the figures FIGURE 1 is an axial sectional view of a launching device and rocket, part of a rifle barrel upon which the launching device is mounted being shown in dot-and-dash lines;

FIGURE 2 is a diagrammatic view illustrating the trajectory of the rocket from launch;

FIGURE 3A is an enlarged sectional view of the rocket taken cross-sectionally through the propellant chamber;

FIGURE 3B is an enlarged sectional view of the rocket taken longitudinally through the propellant chamber;

FIGURE 4A is a schematic frontal view of the rocket grain, indicating the manner in which the exposed area of the grain reduces; and

FIGURE 4B is a schematic side elevational view of the rocket grain and indicating the manner in which the configuration of the grain changes as it is consumed.

Brief general description of invention (refer to FIGURES 1 and 2) This invention is a new and unique method of providing for the transport of warheads, or other payloads, on accurate, essentially straight line trajectories to a preselected target. The method consists of a special design and utilization of solid rocket propulsion to accomplish the desired results in combination with other physical principles and designs. The method can best be described by consideration of the following physical principles that are employed in combination to produce the desired result:

(1) The use of prespin during aiming of the weapon launcher and residual spin during flight. This provides gyro stabilization of the thrust axis of the rocket from the instant of launch to target impact and prevents thrust misalignment about the center of gravity from creating significant trajectory dispersion.

3,245,350 Patented Apr. 12, 1966 (2) A spherical external shape, coupled with an essentially fixed center-of-gravity position are employed to avoid significant aerodynamic precession torques during flight. Another important feature of a spherical shape is that it provides a better mass distribution for gyro stabilization, as contrasted to a conventional cylindrical rocket shape.

(3) The use of rocket thrust forces during the entire or major portion of the flight trajectory. This feature, when properly implemented by maintaining a constant rocket thrust-to-weight ratio, provides the forces to essentially prevent aerodynamic and gravity forces from seriously altering the initial flight path trajectory. This im portant characteristic permits the use of a straightline trajectory so that the range to the target is not a serious factor in weapon accuracy and minimizes the dispersion due to unpredictable atmospheric forces.

The most important feature of the invention consists of the design of the rocket solid propellant geometry and associated thrust history such that the rocket thrust-t0- weight ratio during the powered flight of the rocket is essentially constant (nominally :1 percent or better). This requires that the rocket thrust be constantly reduced, in the proper manner, during burning, by the inherent design characteristics. In addition, the design of the solid propellant and missile hardware must be such that the center-of-gravity of the missile does not vary significantly from the geometric center of the missile during the flight (typically :.01 inch).

Since the thrust-to-weight ratio of the device is essentially constant during burning, and significantly greater than unity (at least 10:1) the force of gravity will not seriously alter the fiight path trajectory. In addition, the atmospheric drag forces in the absence of a crosswind cannot alter the flight path trajectory since they act parallel to the flight path. The crosswind components of drag due to normal surface winds (less than 30 knots) will not seriously alter the trajectory. The small effects that do exist due to crosswinds can be compensated to a suflicient extent by offset aiming.

To illustrate, but not limit the application of the device, the cross-sectional drawing of FIGURE 1 is presented. As described in this drawing, and further in FIGURE 2, it may prove desirable to launch the device from the muzzle of a conventional rifle barrel 1, using an adapter composed of a fixed member 2 which is inserted into the rifle muzzle 1 and is connected to a rotating member 7 in the conventional manner by means of bearings 6. A blank rifle cartridge (not shown) could be used to generate high pressure gases to actuate the translational motion of the friction igniter pin 3, which in the quiescent position is restrained by a preloaded spring 5. Upon activation, the friction igniter pin 3, due to the gas pressure forces from the blank cartridge exceeding the spring restraining forces, moves forward (away from the trifle muzzle). The forward end of the friction igniter pin 3 contacts, by sliding friction, the pyrotechnic ignition pellet 15 in a manner such as to generate suflicient heat to cause ignition of the ignition pellet 15 in turn ignite the pyrotechnic booster pellet 16. The heat generated by the booster pellet 16 is suflicient to ignite the solid propellant rocket grain 18 in the conventional manner. When after activation, and suflicient forward travel of the friction igniter pin 3, the high pressure gases generated by the blank cartridge are expelled to the atmosphere through the exhaust ports 4. The forces from the spring 5 acting against the friction igniter pin are sufficient to return the friction igniter pin 3 to its original (quiescent) position.

The ignition pellet 15 and booster ignition pellet 16 are converted to essentially a gaseous state after burning. By means of the design arrangement, the initial flow of high pressure gas, generated by combustion of the solid propellant rocket grain, is ducted to tangential exhaust ports 8 located on the periphery of the rotating member of the launching adapter 7. The reaction forces created by the rocket gases exhausting through these tangential ports 8 are sufficient to create adequate rotational speeds (l revolutions per second) of the rotating member of the adapter 7 and included payload, during the prelaunch phase of operation. When the desired prelaunch spin rate (10 revolutions per second) of the rotational portion of the adapter 7 and included payload has been attained, the mass of the bobweight 9 and associated centrifugal acceleration provide a force which is suflicient to overcome the preloaded restraining force of the latchspring 10. This action causes the release of the rocket device restraining latch 12 which operates about a pivot '11. The large thrust-to-weight ratio of the rocket device is sufiicient to launch the device after release of the restraining latch 12.

The rocket device which is launched as previously described, consists of a rocket nozzle block which incorporates a conventional nozzle opening 13; a block-to-rocket case seal 14 which prevents the leakage of high pressure gas; a typical solid propellant rocket grain IS; a rocket propellant case 17; a payload explosive charge 19; an external case 20; and a fuzing mechanism 21 which is appropriate for the intended usage.

The sketch of FIGURE 2 illustrates the weapon aiming technique. Here the payload spin axis 22 orientation is shown inclined to the flight path trajectory in conjunction with the major portion of the launching mechanism consisting of the rifle barrel 1, the fixed section of the launcher adapter 2, the rotating member of the launcher adapter 7 and rocket device with its associated payload 23.

The device and associated payload can be scaled in size, weight, and range for a particular application without changing the nature of the invention. For the purpose of illustration, a typical infantry weapon application of the device would consist of an external diameter of three inches. The payload weight could be as large as two pounds in this configuration with an associated straight line range in excess of six hundred yards.

Detailed description of invention FIGURE 3 describes a typical embodiment of the rocket device invention. The prespin and gyro stabilized aspects of the process coupled with the spherical external shape have been adequately described in the previous discussions, therefore, the detailed description will be concerned with the method of obtaining a constant thrustto-weight ratio during burning.

The rocket device as illustrated in FIGURE 3 consists of a rocket nozzle block 13 which includes an appropriate nozzle opening and cut-out to receive the rotating member of the launch adapter and a detent which operates in conjunction with the restraining latch. This nozzle block can be fabricated from a suitable material such as stainless steel and is attached by means of a conventional machine screw thread to the rocket propellant case 17 in a fashion to compress a seal or gasket 14 to prevent the escape of high pressure gas to the atmosphere at this juncture. The end-burning solid propellant rocket grain 18 is ignited by a chain reaction initiating from friction heat acting on the pyrotechnic ignition pellet 15 and spreading by direct contact to the booster pellet 16 as discussed previously. The design of the solid propellant rocket grain is configured in a unique manner such that the thrust-to-weight ratio of the device during burning is constant. The propelling portion (rocket) of the device can be enclosed by a suitable explosive material 19 and external case 20 for use in the intended mission. The increment of distance 24 along the axis of symmetry of the device designates the most fore and aft center-of-gravity positions during burning of the solid propellant. This distance would be held to a minimum and would depend on the particular payload design parameters and maintained within acceptable limits.

The design notation of FIGURE 4 is utilized to describe the means by which the constant thrust-to-weight ratio of the device during burning can be obtained. A common equation that was derived from the physical laws governing the action of rockets, and describes their weight variation with time of burning, isdescribed below:

z -exa where:

W =The initial weight of the rocket (pounds) W=Weight at a time (t) after ignition (pounds) T :Thrust force of the rocket (pounds) e:A standard mathematical constant t=Time after ignition (seconds) I :Rocket propellant specific impulse (seconds) Equation 1 To obtain a desired result from the rocket device, the thrust-to-weight ratio (T/ W) is to be held constant throughout the entire burning period. This can be accomplished if the thrust history is designed in'the following fashion:

Consider that the (T/ W) ratio designated as (n) is to be held constant during burning, Equation 1 then reduces to:

Equation 2 where Equation 3 where:

d =Rate of change of weight (pounds/second) A =Propellant cross-sectional area (square inches) l= Length of propellant along axis of symmetry (inches) dl Burning rate along axis of symmetric (inches/second) V=Z:1 considered to be a constant By combining Equations 3 and 4, the following expression is obtained:

W 'K kt Equation 4 Since in Equation 5, V, U, W and K are considered to be constants, the following expression can be derived:

Equation 5 Equation 6 is a constant (V). The desired relationship can be derived as follows (refer to FIGURE 4):

where where l =initial length of solid propellant (inches) 1: l0 Vt therefore:

IQ "l V Substituting the above expression in Equation 6 yields:

A=Ce -e Equation 7 To illustrate the effects of the rocket design parameters, the constant C, and K are replaced by their equivalents which gives:

W) 1 (a) e 6 IVu where, as before:

Equation 8 It is useful to consider the relationship of Equation 8 when 1:1 This permits a substitution of equivalents to obtain an expression which is more appropriate for the design objective, as follows: when Equation 8 reduces to:

and:

A IVu 0 therefore:

(n lo AoIVul M) 6 IV =6 W IV =8 We in like manner:

e I =6 We in conclusion therefore:

m an A=(% ::)e -e The above equation expresses the necessary propellant cross-sectional variation with length to provide a constant thrust-to-weight ratio of the rocket device during the propellant burning period. It should be noted that the only factors which affect the attainment of this constant relationship of 'thrust-to-weight in the rocket design are A u, and W which are easily controlled to high precision.

I claim:

In combination: a rocket having an outer casing with a substantially spherical external configuration for balancing all aerodynamic forces; a propellant case secured within the outer casing, and having means defining a nozzle opening at the exterior of the outer casing; said propellant case having interior wall means for containing a solid propellant; a solid propellant in the wall means, and having an area exposed in the propellant case to define a boundary of a rocket chamber; said wall means and said propellant having a configuration to produce a substantially constant thrust to weight ratio as the propellant is consumed; a launching device; means supporting the launching device for rotation about an axis; means releasably coupling the launching device and the rocket so that the axis of said nozzle coincides with (the axis of said launching device whereby said rocket rotates wit-h said launching device; and means operative upon predetermined rotation of said launching device for releasing said rocke-t; thereby producing a substantially straight line free-flight trajectory for said rocket.

References Cited by the Examiner UNITED STATES PATENTS 2,391,865 1/1946 Chandler 102-49 2,701,984 2/1955 Terce 891.7 2,775,163 12/ 1956 Vegren 891.7 2,939,449 6/1960 Kortick 42-1 3,045,596 7/1962 Rae 102-50 3,064,381 11/1962 Vilbajo 421 3,088,273 6/ 1963 Adelman et a1. 10249 BENJAMIN A. BORCHELT, Primary Examiner.

SAMUEL FEINBERG, Examiner.

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