US3358603A

US3358603A - Ultra-sonic self-propelled projectile having high l/d ratio

Info

Publication number: US3358603A
Authority: US
Publication date: 1967-12-19
Anticipated expiration: 1984-12-19

Description

W. W. HOHENNER ULTRASONIC SELF-PROPELLED PROJECTILE Dec. 19, 1967 HAVING HIGH L-/D RATIO 3 Sheets-Sheet 1 Filed Dec. 10, 1964 INVENTOR Werner W Hohenne ATTORNEY 1967 w. w. HOHENNER 3,35

ULTRA-SONIC SELF-PROIELLED PROJECTILE HAVING HlGH L/D RATIO Filed D80. 10, 1964 5 Sheets-Sheet I5 United States Patent Ofiice 3,358,003 Patented Dec. 19, 1967 3 358,603 ULTRA-SONIC SELF PROPELLED PROJECTILE HAVING HIGH L/d RATIO Werner W. Hohenner, Adelphi, Md, assignor to Westinghouse Electric Corporation, Pittsburgh, Pa., 21 corporation of Pennsylvania Filed Dec. 10, 1964, Ser. No. 417,322 4 Claims. (Cl. 102-493) This invention relates to a self-propelled projectile.

The primary object of this invention is to provide a new type of projectile which combines extremely lightweight with extremely high velocity to provide a novel projectile that can be self-propelled, launched and aimed with an expendable launching tube. The device utilizes an expandable motor which provides the necessary stabilizing rotational movement before it gets up to speed as well as producing the necessary thrust for the projectile.

The lightness of the projectile is such that it is feasible to produce a net transfer of the energy of the propellant into kinetic energy of the projectile at its effective range so that the mass of the latter serves as its own warhead. The kinetic energy available upon impact is such that if converted to heat by a resisting target, the efiect would be that usually produced by the conventional warhead. This eliminates the necessity of a warhead and reduces by that means the requirements for the weapons system.

It can be said that the present invention departs from the known prior art by a quantitative difference which produces a qualitative difference in results and effectiveness of the system. The velocity of the missile is such that when used in atmospheric environment, if the projectile missed the target, it will automotically become self-destructive.

The expand-able rocket propulsion motor is so designed that it produces continuous thrust from launch to its maximum velocity which is reached after a burning time of less than one second. At bum-out the propulsion ceases. The drag in the air has to be overcome by the momentum. The air drag is very high at high speeds and can be calculated from the well known Equation 1 discussed hereinafter.

Assuming proper target tracking, it is apparent that there is a greater probability of a hit on a target the shorter time interval between firing and the arrival of the projectile at the range of the target. The present invention provides a small caliber, lightweight, single stage inexpensive projectile which is capable of attaining velocities of more than 9,000 feet per second with an initial launching weight of about tWo pounds. In conventional automatic firing weapons that launch projectiles, such as bullets or explosive shells, either the projectile itself has such great weight as to carry impact power or it additionally carries a Warhead which can explode either on impact or in some other controller manner to destroy the target.

In weapon systems having automatic aiming devices, the individual missiles are fired sequentially, the launching mechanism being automatically aimed under control of computers which predict the most probable position of the launching mechanism to cause the projectile to hit the target at the next firing. The computing mechanism can never be 100% accurate in aiming the launching mechanism since the computer must attempt to predict where the target will be at some future instance based on the history of the action of the target. Accordingly, if the path of the target is changed or departs from its previous history of action there is always a certain time lag in correcting the aim. In the meantime the target path may have changed again. This means that there is a certain probability of error that can be symbolized by a volume of physical space which surrounds the target and moves with it. The size of this volume of probable error is a function of the performance of the target and of the accuracy of the fire control system. Because of the lightweight of the projectile of the present invention, a suitable automatically aimed launcher can fire enough of the inexpensive projectiles to fill the cross sectional area of the probable error space almost instanetanously as distinguished from conventional systems in which the projectiles are launched sequentially. The launching of the projectiles in accordance with the present invention may be likened to a shotgun eifect as distinguished from the firing of projectiles in sequence from a rifle.

In brief, the velocity, the lightweight and the effective ness of the projectile of this invention is a result of a novel concept of relating the L/d ratio, that is, the ratio of the length of the combustion chamber to the diameter of the projectile to the rate of combustion in the expendable single stage motor so that a low empty weight and a high mass ratio, M, defined as the ratio of the full weight to the empty weight of the projectile, is accomplished. The rate of combuston is such as to burn the amount of propellant determined by the mass ratio in the shortest possible time without rupturing the combustion chamber made of the material which enters into the calculation of the mass ratio. A projectile made in accordance with the present invention has an L/d ratio within the range of 20 to 50, a mass ratio, M, of approximately 4.1 and a final speed of approximately 10,000 ft./sec. The L/d ratio and the mass ratio, M, are very closely related and are the primary parameters which determine the capabilities of the projectile and make it possible to fabricate the projectile from thin-walled mass produced tubing. The burning rate is controlled by a combustion inhibitor to determine the combustion interval and prevent the development of eX plosive pressures in the combustion chamber.

An important feature of the invention is that its extreme lightweight and size makes it possible to use lightweight launchin-g tubes which may be made of such materials as cardboard or plastics. It is, of course, necessary to protect the ends of the projectile during storage and transportation and therefore some type of packaging unit is necessary. The same protecting tubes that serve as the packaging units can also be used as the launching tube. Because of the light weight of the launching tubes and the missiles, a multiple of these packaging tubes carrying projectiles can be assembled in the nature of a honeycomb array and be carried by a dirlgible mount which can be aimed manually or by suitable automatic aiming apparatus at very high angular rates. This makes it possible to provide a launching mechanism suitable for mass firing of these projectiles with great accuracy.

In the present invention the jet of gases from the reaction motor produces a stabilizing spin of the projectile about the longitudinal axis before it leaves the launching tube. This is a very important feature in stabilizing the flight of the projectile since by the very nature of the present invention the casing of the projectile may be made of tubing made by mass production methods and therefore may not be dynamically balanced within close tolerances. The rotational spin of the projectile is produced by small turbine fins which may be produced by stamping operations and which may contribute to some dynamic unbalance. The spinning of the projectile as soon as it starts from the launching tube is also effective in preventing small imperfections in the Ventun' from causing excessive wobbling of the projectile.

The speed of the present projectile will reach its maximum value about one-half second after the instant of launching. The maximum velocity will be approximately 9,000 feet per second, or more, and its average speed will be about 5,000 feet per second. Although it is considered that the present invention is primarily a surface-to-air projectile for short ranges, calculations indicate that the present projectile is capable of penetrating about three inches of armor plate because of the high velocity obtained and therefore the projectile makes a good surfaceto-surface weapon capable of destroying tanks and other armored vehicles within the range of the projectile. Because of the lightweight of the device and because the projectile can be fired from a launching tube of extreme lightweight it can readily be fired by an infantry man who can also carry. a reasonable supply of projectiles.

The above objects, stated in terms of the capabilities and summarized features of the invention, together with other objects and advantages of the invention will best be understood from the following description when. taken in connection with the accompanying drawings, in which:

FIGURE 1 is a side elevation view, approximately to scale, of a self-propelled projectile in accordance with this invention;

FIG. 2 is an enlarged partial sectional elevation view of FIG. 1;

FIG. 3- is an end view of FIG. 2;

FIGS. 3A, 3B, 3C and 3D are transverse sectional views at a selected plane indicated on FIG. 3, illustrating successive stages of the controlled combustion of the solid propellant;

FIG. 4 is a graph illustrating the comparison between the velocity-time history of a projectile in accordance with the present invention and a projectile fired from a contemporary. 30 millimeter cannon;

FIG. 5. is a distance-time graph for a projectile in accordance with the present invention as compared to the projectile from a contemporary 30 millimeter cannon;

FIG. 6 is an isometric view of a' launching module and packaging case for the present invention;

FIG. 7 is a side elevational view of a dirigible launching mechanism for the present invention; and

FIG. 8 is a front elevational view of FIG. 5.

An embodiment of the invention, illustrated in FIG. 1, comprises a lightweight tubular body 10 having a front conical section 11 and an aft section 12 shaped to constitute aVenturi and terminating in a frustoconical' portion 12a. The body 10 is preferably made of thin metal alloy tubing, similar to that used as electrical conduit and having a wall thickness of approximately .035 inch. One of the important aspects of the present invention is that the present concept makes possible the utilization. of a major component that is otherwise widely used and is therefore. relatively inexpensive because of the large quantities in which it is produced by mass production methods.

The pointed front conical section 11 is preferably made of heat resistant lightweight metal or plastic, such as nylon or Teflon, and its primary purpose is to give the necessary aerodynamic properties to reduce air drag. The exact manner in which the conical section 11 is attached to the main body portion 10 constitutes no part of the present invention. It may be aifixed to the body section 10 in accordance with manufacturing techniques well known. In the illustrated embodiment the front section' of the body 10 is provided with a frusto-con-ical portion 13 which is an extension of the body 10. A squib 14 may be mounted by a press fit into a serrated bore recess in the section 11. The squib 14 may threadedly engage a cylindrical forward extension of the section 13. Suitable conductors 15a and 15b, connected to the squib, extend through the walls of the section 11 to provide means through which the squib may be electrically ignited to start the combustion of a core 16 of solid propellant, the outside diameter of which fits the inside of the cylindrical section 10. The core 16 conventionally has a tapered bore 17 which increases in size from front to rear. The bore 17 constitutes the venting channel through which the combustion gases flow to the Venturi section 12.

The aft section 12 is provided with a plurality of vanes 12b constructed and feathered in conventional manner so that the action of the gas jet from the Venturi will cause the projectile to spin about its longitudinal axis. The spinning will be initiated as soon as combustion begins and before the projectile leaves the launching tube.

Heretofore, it has been considered impossible to reduce to about 0.5 second the burning interval of a propellant in a single stage reaction motor where the ratio of the length of the combustion chamber to its cross sectional dimension has the values in accordance with this invention. The ordinary solid fuels for ordinary rocket propulsion is merely modified explosive material. In producing explosive bombs and explosive charges it is desired to have the fastest possible burning rate so that the highest possible pressures can be developed. On the other hand, in rocket propulsion 'the requirements are quite the opposite in that it is desired to produce a controlled rate of burning in order to produce a sustained thrust over a longer period of time and this means that it is desired in projectile propulsion to limit the maximum pressure but to produce a sustained thrust over a longer burning interval.

The end results of the present invention are determined by the novel L/ d ratio, its relation to the mass number on the basis of state-of-the-art solid fuels and the controlled burning rate. The novel proportioning of the factors of the. present invention is determined on the basis of the air drag which can be stated mathematically, as follows:

C is the drag coefficient; and g is acceleration of the earth in ft./sec.

From Equation 1 it is obvious that air drag, on a projectile depends upon two parameters which are accessible to modification at will, within limits, namely, the cross section, A, and the drag coefiicient C The drag coefiici'ent, can be minimized by selecting a favorable aerodynamic form.

It is apparent from Equation 1 that the primary parameters affecting drag are the area, A, of the projectile, which is a function of the square of the diameter, and the velocity. Therefore, by keeping the cross sectional area low the velocity can be increased substantially without increasing the drag. As the diameter is reduced to reduce the cross sectional area, A, there develops a problem of controlled burning of the propellant and venting the gases to develop the desired thrust without high explosive pressures that would necessitate heavy wall thickness of the projectile which would then defeat the basic objective of the invention, namely, a high mass ratio.

By choosing a high L/d ratio in accordance with the present invention of a value at least twice that conventionally used, in the range from 20 to 50, it is necessary to provide a selected amount of propellant and controlled rate of combustion in such a manner that a final velocity of 9,000 to 10,000 ft./sec. would be obtained in a time interval in the range between 0.5 and 1.0 second without at the same time reaching a pressure in the combustion chamber which would rupture the body 10. FIGS. 6 and 7 graphically illustrate the performance characteristics of the present novel projectile as compared to that of a contemporary 30 mm. cannon in terms of velocity and distance attained as a function of time.

The amount of propellant of a known characteristic required to propel a projectile to a certain maximum speed may be determined from the well known equation AV=I ,-g.log, M 2

where AV is the speed change (ft/sec.) I is the specific impulse lb. thrust lb./sec.

g is the gravity constant (32.2 ft./sec. and M is the mass ratio (a propellant characteristic in AV 10,000 g Then If the weight empty is assumed to be 1 unit, the weight full must be 4.1 units. Accordingly, the amount of propellant must be run W =4.11=3.1 weight units (5 The above calculations give a good median value to use in arriving at final proportioning of factors for providing the novel self-propelled projectile in accordance with this invention; which would have the desired mass ratio, M, and the desired burning rate without developing explosive rupturing pressures in the combustion chamber.

To control the rate of combustion and thus prevent the development of rupturing pressures in the combustion chamber, suitable combustion inhibitor means are provided on the inner surface of the propellant core 16. To this end, the core 16 has a plurality, in the illustration, two, substantially U-shaped combustion inhibitor shields 18 and 19 secured to and imbedded therein. Each of the inhibitor shields have arcuate portions 18a and 19a, respectively, fitting the inside of the bore 17 and parallel flanges 18b and 1%, respectively, extending generally rad-ially of, and imbedded in the core 16. These shields are made of suitable heat reflecting material capable of preventing of initiation of combustion on that portion of the inner surface of the bore 17 against which the arcuate portions are engaged. Also the flanges prevent initiation of combustion of portions of the core constituting ribs 16a and 16b protected by the flanges 18b and 1%, respectively.

This is illustrated in FIGS. 3A and 3D, inclusive, where FIG. 3A represents the condition before combustion is initiated and the other figures represent the condition at the same section at successive intervals after combustion starts. When the squib 14 is ignited, almost instantaneously combustion is initiated on all of the exposed inner surface of the bore 17. FIG. 33 represents a condition that exists throughout the length of the bore 17 at an instant immediately after combustion starts. The surface 11npro tected by the shields 18 and 19 will immediately burn with the burning surface progressing radially leaving the ribs 16a and 16b. As the combustion continues with the gases being spewed into the gradually enlarging central opening combustion will begin to undercut the ribs as at 16c and 16d, as illustrated in FIG. 3C. Preferably the dimensions of the flanges and the arcuate portions of the combustion inhibitor shields are so related as to cause the ratio of the area of the burning surface to the area of the venting channel defined by the burning surface to remain substantially constant during the burning interval. As combustion proceeds the ribs will be completely undercut and the shields 18 and 19 will be completely vaporized and the venting channel will assume a general circular cross section as indicated in FIG. 3D, with any remnants of the ribs finally being completely consumed.

In the illustrated embodiment, all of the burning takes place in a time interval ranging from 0.5 to 1.0 second.

The shields 18 and 19 are preferably made of a metal having a low melting point but having a high coeflicient of reflectivity for heat rays so as to prevent burning under the shield until the ribs 16a and 16b have been undercut. The effect of the shields is to increase the surface area over which the combustion zone must creep, thereby increasing the burning interval and reducing the maximum pressure developed in the combustion chamber. The shields 18 and 19 are preferably made of an aluminum foil. Although only two combustion inhibitor shields are shown for illustrative purposes it will be apparent that the number is not critical and also that the arrangement of such shields may be changed to carry out the objective of the invention.

It is to be understood that the mass ration, M, is dependent upon the characteristics of contemporary materials, such as propellants and the material of the tube 10. The maximum permissible internal pressure in the combustion chamber is dependent upon the tensile strength of the walls of the combustion chamber, that is, the tube 10. It will be seen therefore, that it is within the contemplation of this invention that the mass ratio, M, may be made even higher than previously indicated, and therefore the performance of the projectile improved, if the tube 10 is made of a material in which the ratio of the tensile strength to its weight is higher than for the material indicated for the illustrated embodiment. Good performance can be obtained if the mass ratio is at least substantially 2.5 or above.

It has been mentioned previously that the self-propelled projectile of the present invention is of such small size and light weight as to be adaptable to simultaneous multiple launching, either manually or automatically. In FIG. 6 there is illustrated a group of launching tubes 25 each of which carries a projectile in accordance with this invention. The group of launching tubes is held in assembled relation by suitable retaining bands 26. The assembled groups constitute a launching module, a number of which may be stored in

magazines

27 and 28 on either side of a dirigible aiming and launching mechanism 29, illustrated in FIGS. 7 and 8. The launching mechanism may be aimed by means of a sighting device 31 manually under the control of a human operator 32. Suitable means, not shown, are provided through which the operator can selectively launch the projectiles. Obviously, the launching mechanism may be aimed automatically, if desired. The launching mechanism constitutes no part of the present invention but is illustrated merely for the purpose of more graphically presenting the adaptations and capabilities of the novel projectile.

It will be readily apparent to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention.

I claim as my invention:

1. A self-propelled projectile comprising a cylindrical tubular element constituting a combustion chamber and generally determining the outer relative proportions of said projectiles, a solid fuel propellant grain in the form of a cylindrical annulus, the inner surface of which forms a central passage for the passage of combustion gases for producing thrust for said projectile, said chamber having a ratio of length to diameter greater than 20, said projectile having a mass ratio greater than 2.5 and means for controlling the combustion rate of said propellant grain so that the latter is burned at such a rate that combustion is completed in a time interval of 1.0 second or less to thereby keep the maximum pressure in said combustion chamber within tolerable limits of stress in the walls of said combustion chamber set by the mass ratio.

2. A self-propelled projectile comprising a cylindrical tubular element constituting a combustion chamber, a solid fuel propellant grain in the form of a cylindrical annulus, the inner surface of which forms a central passage for the passage of combustion gases for producing thrust for said projectile, said chamber having a ratio of length to diameter grater than 20, said projectile having a mass ratio greater than 2.5 and means for controlling the combustion rate of said propellant, grain so that the latter is burned at such a rate that combustion is completed in a time interval of 1.0 second or less so as not to produce maximum pressure in said combustion chamber above the tolerable limits of stresses in the combustion chamber Walls set by the mass ratio.

3. A self-propelled projectile comprising a cylindrical tubular element constituting a combustion chamber, a solid fuel propellant grain in the form of a cylindrical annulus, the inner surface of which forms a central passage from the front to the back of said projectile for combustion gases for producing thrust for said projectile, said chamber having a ratio of length to diameter greater than 20, said projectile having a mass ratio greater than 2.5 and means for controlling the combustion rate of said propellant grain so that the latter is burned at such a rate that all combustion takes place in a time interval of 1.0 seconds or less.

4. A self-propelled projectile comprising a cylindrical tubular element constituting a combustion chamber and the support structure for said projectile, a solid fuel propellant grain in the form of a cylindrical annulus, the inner surface of which forms a central passage substantially from end-to-end of said projectile for combustion gases for producing thrust for said projectile, said chamber having a ratio of length to diameter greater than 20, said projectile having a'mass ratio greater than 2.5, said solid fuel propellant grain having combustion inhibitor means in the form of heat reflecting foil covering a portion of the inner surface of the central passage and having flange portions extending into said propellant grain to delay the combustion of certain portions of said propellant so that the ratio of burning surface to the area of said passage remains substantially constant during the burning interval, said means being so adjusted that the combustion is completed in a time interval of one second or less.

References Cited UNITED STATES PATENTS BENJAMIN A. BORCHELT, Primary Examiner.

SAMUEL FEINBERG, Examiner.

V. R. PENDEGRASS, Assistant Examiner.

Claims

1. A SELF-PROPELLED PROJECTILE COMPRISING A CYLINDRICAL TUBULAR ELEMENT CONSTITUTING A COMBUSTION CHAMBER AND GENERALLY DETERMINING THE OUTER RELATIVE PROPORTIONS OF SAID PROJECTILES, A SOLID FUEL PROPELLANT GRAIN IN THE FORM OF A CYLINDRICAL ANNULUS, THE INNER SURFACE OF WHICH FORMS A CENTRAL PASSAGE FOR THE PASSAGE OF COMBUSTION GASES FOR PRODUCING THRUST FOR SAID PROJECTILE, SAID CHAMBER HAVING A RATIO OF LENGTH OF DIAMETER GREATER THAN 20, SAID PROJECTILE HAVING A MASS RATIO GREATER THAN 2.5 AND MEANS FOR CONTROLLING THE COMBUSTION RATE OF SAID PROPELLANT GRAIN SO THAT THE LATTER IS BURNED AT SUCH A RATE THE COMBUSTION IS COMPLETED IN A TIME INTERVAL OF 1.0 SECOND OR LESS TO THEREBY KEEP THE MAXIMUM PRESSURE IN SAID COMBUSTION CHAMBER WITHIN TOLERABLE LIMITS OF STRESS IN THE WALLS OF SAID COMBUSTION CHAMBER SET BY THE MASS RATIO.