US3319566A - Non-spin rockets and their guidance - Google Patents

Description

May 16, 1967 P. v. CHOATE ETAL NON-SPIN ROCKETS AND THEIR GUIDANCE 2 Sheets Sheet 1 Filed July 25, 1966 I fi :1 Q. mum W m mum d b3 & kmn m 2|... J a mmh 1M1 WY 5w mm (I t w on r a g? Q 3 wm I A a av! C \V MN M m y Q t s I... Q 0 O M mm mm PAUL v.

A77URNE Y May 16, 1967 P. V. CHOATE ETAL NON-SPIN ROCKETS AND THEIR GUIDANCE Filed July 25, 1966 FIGS FIG. IO

FIG.

2 Sheets-Sheet 2 INVENTORS PAUL V. CHOATE MICHAEL A. NEE

ATTORNEY United States Patent Ofiiice 3,319,566 NQN-SPIN ROCKETS AND THEIR GUHDANCE Paul V. Choate, Milton, and Michael A. Nee, Melrose, Mass, assignors to Norris Industries, 1116., Los Angeles, (Zalih, a corporation of California Filed July 25, 1966, Ser. No. 570,124 10 Claims. (Cl. 1tl2-49.7)

The present invention relates to non-spin rockets and to their guidance, the rockets being of any size ranging from those that are pad-based to small rockets such as .those that may be launched from shoulder-supported launchers. The present invention is a continuation-inpart. of our copending application, Ser. No. 446,275, filed Apr. 7, 1965, and now abandoned.

While the invention is applicable to the entire range of non-spin rockets, particular reference is made herein to rockets that are used as missiles of the type that do not have sophisticated guidance systems. The novel features and advantages of the invention may, furthermore, be best illustrated in connection with rockets capable of being fired from a shoulder-supported launcher. Such rockets have proved to be effective even though they are essentially single shot weapons, making first round accuracy of great importance. While many considerations are involved in seeking improved accuracy, stability in flight is a most important factor and is aliected by the manner in which the thrust of the resulting gas stream is applied. Conventional rockets have their nozzles axially of their rear ends. Any eccentricity of the gas stream forces is potentially capable of throwing such a rocket oti course since it provides long moment arm advantage that is responsive to relatively small eccentric forces of the propelling gas stream.

Rockets, like other projectiles, require stabilization and, except for the present invention and apart from guidance systems, there are two known methods by which stabilization can be eflected. The two methods, as set forth in ORDP 20-246, Ordnance Corps Pamphlet, Ordnance Engineering Design Handbook, Artillery Ammunition Series, Section 3, Design for Control of Flight Characteristics, are spin stabilization and fin stabilization. If a projectile is spin stabilized, any increase in its length increases the difliculty of or makes it impossible to effect its stabilization. For that reason, there is a diameterto-length relationship that is characteristic of spin rockets,

these lengths being in the order of five to seven times their diameters, see The Exterior Ballistics or Rockets, by Davis, Pollen, and Bletzer. Spin rockets rotate several hundred times a second. Non-spin rockets have hitherto relied upon fins to stabilize their flight and their length-to-diameter ratios are greater than those of rockets requiring spin stabilization. As far as stability in flight is concerned, there are no limits or the extent to which these ratios may be increased with over all length being determined by such factors as launcher limitations and ease of handling.

A primary objective of the present invention is, accordingly, to provide rockets of increased stability in flight by a third basis of stabilization and this objective is attained by providing a rocket having a motor with the rocket having an elongate-d, rearwardly extending portion and the motor having a rearwardly opening nozzle surrounding the rocket approximately in the zone of the center of gravity of the rocket and the rearward rocket portion being dimensioned to be enveloped by the expanding gas stream from the nozzle and, being free of surface portions that would so interrupt the gas stream as to afiect adversely its holding action as would be the case if the rocket were spun.

With such a rocket, guidance in accordance with the invention is achieved, the guidance requiring not only estates Patented May is, rear that an annular rearward stream of gas from the rocket motor be provided that is coaxial with the rocket axis but also that the accelerating forces of the gas stream be applied to the rocket in an annular zone approximately inclusive of the rockets center of gravity and that the gas stream envelope its rear portion with the inward components of the forces resulting from the expansion of the gas stream applied thereto to hold the rocket on course.

A rockets accuracy is also affected by its velocity as is its range. Another primary objective of the invention is, accordingly, to provide high velocity rockets even when the rockets are to be fired from shouldersupported launchers. This objective is attained by providing for a rocket in accordance with the invention being launched with a propellant charge, sometimes herein referred to as the flight assisting or sustaining charge, effective when the rocket is air borne to provide the desired velocity, such charge being ignited either within the launcher or when air borne or becoming fully effective when air borne, the rocket being launched at a relatively low velocity as by a charge, sometimes herein referred to as the launching charge, contained in the rocket or within the launcher, or by providing that the flight assisting or sustaining charge be not fully effective until the rocket is air borne.

In accordance with the invention, the opportunity for eccentric gas stream forces to affect a rockets course is not only minimized by the manner in Which the accelerating forces are applied thereto but also such forces are prevented from so doing by the use of the expanding gas stream as a gas envelope holding the rocket on course with a long moment arm advantage.

The use of tail fins is desirable when the fuel supply burns so quickly that the thrust is applied to the rocket during only a minor part of its flight. Where the motor is so designed as to provide thrust throughout a predetermined range of a rocket, fins are not necessary. In that case, however, rotation of that rocket is desirable only as a means for increasing its accuracy by offsetting the efiect of any unevenness in the gas flow through the annular nozzle. Another objective of the invention is, accordingly, to provide a rocket that rotates in flight at a rate far below a rate useful in spin stabilization thus to avoid upsetting the holding action of the gas stream, the rotative forces being applied thereto in the zone of its center of gravity.

The gas stream may be used to effect changes in course in conjunction with various guidance systems such as adjustable tail fins, longitudinal fins in the gas stream that can be turned, thrust deflectors in the nozzle, or by the use of radial gas jets, preferably in the area of the gas envelope.

In the accompanying drawings, there are shown illustrative embodiments of the invention from which these and other of its objectives, novel features, and advantages will be readily apparent.

In the drawings:

FIGURE 1 is a longitudinal section of a rocket in accordance with the invention, the rocket being of a type utilizing both launching and booster or flight sustaining propellant charges,

FIGURE 2 is a fragmentary view of the throat of another motor in accordance with the invention, the throat being defined in part, by an internal plug,

FIGURE 3 is a like view of the throat of another embodiment of the invention illustrating an adjustable internal plug,

FIGURE 4 is a somewhat schematic view of the rocket of FIGURE 1 in flight, illustrating the pressure envelope provided by the gas stream,

.FIGURE 5 is a fragmentary and partly sectioned view illustrating the annular nozzle of another rocket in accordance with the invention,

' FIGURE 6 is a section transversely of the rocket illustrated by FIGURE 5 rearward-1y of its nozzle, FIGURE 7 is a like view illustrating a rocket of the type illustrated by FIGURES l-4 and its launcher,

FIGURE 8 is a similar view illustrating another rocket in accordance with the invention, the rocket also being of a type utilizing both launching and booster or flight sustaining propellant charges, the launching charge being of the external motor type,

FIGURE 9 is a like view of yet another similar rocket in accordance with the invention and its launcher of the recoilless rifle type,

FIGURE 10 is a similar view of another rocket in accordance with the invention and its launcher, the rocket having a single charge, and

FIGURE 11 is a similar view of the rocket whose nozzle is illustrated in FIGURES 5 and 6, the rocket also having a single charge and capable of being launched from a launcher such as that illustrated by FIGURE 10.

The rocket detailed in FIGURE 1 includes a launching motor, a flight sustaining or booster motor, and a war head generally indicated at 10, 11, and 12, respectively.

The launching motor 10 consists of a tubular chamber 13 for a fuel supply 14 provided with an igniter 15 in an indicated ignition circuit 16. The launching motor 10 also has a rearwardly opening, axial nozzle 17. The out side diameter of the chamber 13 is increased as at 18 at the forward end and is interiorly threaded to receive a closure 19. The closure 19 is shown as having a central socket 20 threaded to receive the axial projection 21 of a cap 22. Fins 23 are connected by pivots 24 to ears 25 on the nozzle 17 so that they may be folded forwardly into their dotted line positions shown in FIGURE 1 when the rocket is inserted into its launcher 26, see FIGURE 7. The maximum diameter of the nozzle 17 may be greater than the diameter of the motor 10 being limited only to the bore dimension of the launcher. Springs 27, connected to the fins 23 and to the ears 25, are operative to urge the fins 23 into their erected position, once the rocket is air borne.

The flight sustaining motor 11 consists of a cylindrical, rearwardly opening housing 28 having an axial bore 29 in its front wall 30 threaded to receive the forward end of a post 31 the rear end of which is threaded into a socket 32 in the front face of the cap 22. The post 31 has a bore 33 opening through its front end and a series of radial ports 34 effecting communication between the bore 33' and the interior of the housing 28. The fuel supply of the motor 11 is shown as in the form of a sleeve 35 located within the housing 28 and within the bore 33 there is an igniter 36 in an indicated ignition circuit 37.

The inner surface of the housing 28, at the rear end thereof, is outwardly flared as at 38 and the housing 28 is so dimensioned relative to the front part of the chamber 13 that the annular flared surface 38 overlies the junction of the chamber portion 18 and the rounded edge of the cap 22 to define a generally indicated annular nozzle 39. It is essential to note that the nozzle 39 is approximately in the zone of the center of gravity of the rocket as will be apparent from FIGURES 4 and 7.

The front face of the housing wall 30 has an annulus 40 in support of a sleeve 41 receiving the inner end 42 of the war head 43, the front face 44 of which supports an axial, collapsible nose 54. A faring 46, of the same outside diameter as the housing 28, completes the front part of the rocket.

In practice, the rocket just disclosed is placed in its launcher 26. When the rocket is to be launched, the circuit 16 is closed and the launching motor 10 becomes operative. The motor 11 does not become operative until the rocket is air borne, an effect attained as by providing, for example, that the circuit 37 be closed with the closing of the circuit 16 but that the igniter 36 be relatively slow burning. The motor 10, accordingly, is effective to place the rocket in flight but the principal thrust is derived from the motor 11. Gas from the motor 11 is delivered through the rearwardly disposed, annular nozzle 39 with the thrust being applied annularly and approximately at the center of gravity of the rocket. As will be apparent from FIGURE 4, the propelling stream of expanding gas is in the form of an envelope 47 applying increased pressure against the rear part of the nozzle so that while the rocket is subjected to propelling forces, it is held on course by the gas stream, ensuring both accuracy and long range.

Reference is made to FIGURE 2 in which the motor 10A has its throat defined in part by an internal plug 48 underlying the end 38A of the housing 28A of the motor 11A and also to FIGURE 3 wherein the internal plug 48A is in engagement with threads 49 on the motor 10B to enable its position to be varied relative to the shoulder 38B of the housing 28B of the motor 113.

Rockets of other types may be made in accordance with the invention. For example, the rocket shown in FIG- URE 8 is adapted to be fired from the launcher 50, and has a launching motor, a flight sustaining motor, and a war head generally indicated at 10C, 11C, and 12C, respectively. The war head 12C may be identical to the war head 12 and its motor is identical to the motor 11 except that, in place of the post 31, there is an axial rod 51 having a member 52 defining with the housing 28C the annular nozzle 39C approximately in the zone of the center of gravity of the rocket when in flight. The rod has a second member 53 at its rear end adjacent which are fins 54. The fuel supply of the motor NC is shown in dotted lines and is indicated at 14C, the dotted line showing being necessary because the rocket of FIG- URE 8 is illustrated as air borne. in its launcher 50, the launcher serves as the casing of the motor 10C and defines, with the second member 53, its nozzle.

The rocket of FIGURE 9 is for use with the launcher 55 and, like the rocket of FIGURE 8, its launching motor 10D is completed by its launcher 55 when loaded therein. As the launcher 55 has a nozzle 56 at its rear end, the rod member 51D, while having a member 52D defining an annular nozzle 39D with the housing 28D of the motor 11D, is not provided with a member such as the member 53 but the rocket of FIGURE 9 may otherwise be identical to the rocket of FIGURE 8 and its annular nozzle 39D is approximately in the zone of its center of gravity.

The rocket illustrated by FIGURE 10, like that shown in the co-pending application of Paul V. Choate, Ser. No. 395,242, filed Sept. 9, 1964, is to be launched from the launcher 57 and has a single motor 58 and a war head 12E, the single motor having its annular nozzle 39E approximately in the zone of the center of gravity of the rocket.

Reference is now made to the embodiment of the rocket illustrated by FIGURES 5, 6 and 11. Like the rocket illustrated by FIGURE 10, this rocket can be launched from the launcher 57, has a single motor 59 having its annular nozzle 39F approximately in the zone of the rockets center of gravity and includes a war head 12F. The motor 59, however, is adapted to provide thrust throughout the maximum range of the rocket and the rocket does not have tail fins.

As shown in FIGURES 5 and 6, the nozzle 39F is shown as having three fins 60 spaced apart in its throat and disposed angularly with respect to the ro-ckets axis whereby the rocket is rotated by the gas stream through the nozzle 39F, the rate of rotation being desirably in the vicinity of 15 to 25 r.p.s.

From the foregoing, it will be appreciated that each rocket in accordance with the invention, when in flight,

When such a rocket is has the propelling force of its gas stream applied in an annular zone approximately inclusive of the rockets center of gravity with the gas stream so enveloping its rear portion that the rocket is held on course by the gas stream, the rocket being, rearwardly of its nozzle, wholly free of structure that would so interupt the .gas stream as to have a biasing effect on the rocket.

We claim:

1. In a non-spin stabilized rocket, a motor having a chamber, and a propellant charge in said chamber, said motor including a rearwardly opening annular nozzle having a throat and surrounding said rocket in a zone approximately inclusive of the center of gravity of said rocket during the burning of said charge, said motor including a mixing chamber having rearward, direct, annular and substantially unobstructed communication with the throat of said nozzle, a line lengthwise of any portion of the part of said rocket rearwardly of said nozzle being capable of being presented to the viewer as a fore-and-aft line of the rocket at least when the rocket is in flight, the part of said rocket rearwardly of said nozzle being cylindrical and of a diameter less than the outside diameter of said nozzle in a zone commencing immediately adjacent the nozzle and extending substantially throughout its length, and the nozzle being disposed to direct the propelling gas stream resulting from the burning of said charge substantially parallel to the flight axis of the rocket and close to said rearward part along the entire cylindrical length thereof whereby, during the burning of said charge, the rocket is held on course by a non-biasing, expanding gas stream enveloping said rearward rocket part.

2. The rocket of claim 1 in which the portion of the rocket rearwardly of the nozzle includes a second motor having a chamber, a propellant in the chamber of the second mot-or, and the second motor includes a rearwardly.

5. The rocket of claim 1 in which the rocket includes first and second cylindrical portions arranged one behind the other in axial alinement, each portion has a front end wall and an axial post interconnects the end walls, the rear end of the forward portion being open and overlying the forward end of the rearward cylindrical portion and establishing therewith the rearwardly opening, annular nozzle.

6. The rocket of claim 5 in which the post is hollow and opens through one wall, the forward cylindrical portion is the propellant chamber and the post has ports opening therein.

7. The rocket of claim 5 in which the post is hollow and opens through one wall, ignition means extend within the post through that wall, the forward cylindrical portion is the propellant chamber and the post has ports opening therein.

8. The rocket of claim 5 in which the forward end of the front end wall of the rearward cylindrical portion is of conic form.

9. The rocket of claim 1 and a series of equally spaced guidance fins attached to the rearward part of the rocket rearwardly of the nozzle and, at least in flight, extending lengthwise of the rocket and outwardly to an extent such that, during the burning of the charge, the fins are subjected to the holding action of the gas stream.

10. The rocket of claim 1 in which the motor is located forwardly of the nozzle and the mixing chamber is at the rear of the propellant-containing chamber and forms a part thereof.

References Cited by the Examiner UNITED STATES PATENTS 2,503,271 4/ 1950 Hickman 102-49 2,946,261 7/ 1960 Crockett 102-49 3,176,615 4/ 1965 De Matthew 102-49 BENJAMIN A. BURCH-BLT, Primary Examiner. V. R. PEN-DEGRASS, Assistant Examiner.

Claims (1)

1. IN A NON-SPIN STABILIZED ROCKET, A MOTOR HAVING A CHAMBER, AND A PROPELLANT CHARGE IN SAID CHAMBER, SAID MOTOR INCLUDING A REARWARDLY OPENING ANNULAR NOZZLE HAVING A THROAT AND SURROUNDING SAID ROCKET IN A ZONE APPROXIMATELY INCLUSIVE OF THE CENTER OF GRAVITY OF SAID ROCKET DURING THE BURNING OF SAID CHARGE, SAID MOTOR INCLUDING A MIXING CHAMBER HAVING REARWARD, DIRECT, ANNULAR AND SUBSTANTIALLY UNOBSTRUCTED COMMUNICATION WITH THE THROAT OF SAID NOZZLE, A LINE LENGTHWISE OF ANY PORTION OF THE PART OF SAID ROCKET REARWARDLY OF SAID NOZZLE BEING CAPABLE OF BEING PRESENTED TO THE VIEWER AS A FORE-AND-AFT LINE OF THE ROCKET AT LEAST WHEN THE ROCKET IS IN FLIGHT, THE PART OF SAID ROCKET REARWARDLY OF SAID NOZZLE BEING CYLINDRICAL AND OF A DIAMETER LESS THAN THE OUTSIDE DIAMETER OF SAID NOZZLE IN A ZONE COMMENCING IMMEDIATELY ADJACENT THE NOZZLE AND EXTENDING SUBSTANTIALLY THROUGHOUT ITS LENGTH, AND THE NOZZLE BEING DISPOSED TO DIRECT THE PROPELLING GAS STREAM RESULTING FROM THE BURNING OF SAID CHARGE SUBSTANTIALLY PARALLEL TO THE FLIGHT AXIS OF THE ROCKET AND CLOSE TO SAID REARWARD PART ALONG THE ENTIRE CYLINDRICAL LENGTH THEREOF WHEREBY, DURING THE BURNING OF SAID CHARGE, THE ROCKET IS HELD ON COURSE BY A NON-BIASING, EXPANDING GAS STREAM ENVELOPING SAID REARWARD ROCKET PART.

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