US3144828A

US3144828A - Spin-stabilized rocket projectile

Info

Publication number: US3144828A
Application number: US290288A
Authority: US
Inventors: Wehlow Karl-Otto; Woyt Fritz; Renner Hermann
Original assignee: Rheinmetall GmbH
Current assignee: Rheinmetall Industrie AG
Priority date: 1962-06-30
Filing date: 1963-06-21
Publication date: 1964-08-18
Anticipated expiration: 1981-08-18
Also published as: DK112497B; CH400839A; BE634186A; GB1014847A; DE1213760B

Description

3 Sheets-Sheet l KARL-OTTO WEHLOW ET AL SPIN-STABILIZED ROCKET FROJECTILE dflllrllllllllllfillllIlllfll Filed June 21, 1965 L, M .mdf eww @Wyk D .m M 0 f f n t a 0.35/ )www By /M/f/iak Allg. 18, 1954 KARL-OTTO wEHLow ETAI. 3,144,828

SPIN-STABILZED ROCKET PROJECTILE 3 Sheets-Sheet 2 Filed June 2l, 1965 5 H jm@ Aug. 18, 1964 KARL-OTTO wEHLow ETAL 3,144,823

SPIN-STABILIZED ROCKET PROJECTILE B), Hel-17711112 rma/- United States Patent O 3,144,828 SllN-STABKLIZED ROCKET PRXECTHE Karl-@tto Wehlow, Fritz Woyt, and Hermann Renner,

Dusseldorf, Germany, assignors to Firma Rheinmetall Grmbl-I., Dusseldorf, Germany f Filed .inne 21, 1963, Ser. No. 290,288 Claims priority, application Germany ,lune 30, 1962 4 Claims. (Ci. HB2- 49) This invention relates to spin-stabilized rocket projectiles. By comparison with fin-stabilized rocket projectiles, spin-stabilized rocket projectiles have the important advantage of greater accuracy, since the properties of the gyroscope ensure greater insensitivity to inaccuracies of manufacture and irregularities of gas ilow in the combustion chamber and thrust nozzle of the rocket and any disturbing influence introduced by a tail unit is eliminated.

A substantial advantage for military use is the simplicity of construction obtained by the omission of the generally fragile tail unit. Omission of the tail unit improves the transportability of the rocket and makes handling easier. In addition, a simple ring device can be used.

However, when the rocket has to contain a hollow charge projectile, for example for fighting tanks, the high speed required for satisfactory gyroscopic or spin stabilization is a drawback, since all the hollow-charge projectiles known heretofore exhibit a marked drop in pene trating power when a certain speed fof revolution is excede-ed.

The stability number of a spin-stabilized projectile is proportional to the square of the axial moment of inertia and the angular velocity, and is in inverse proportion to the transverse moment of inertia and the moment of resultant air force about the transverse axis ofV the projectile, which tends to turn the projectile out of its ightpath position and thereby initiate a precession movement.

The stability number a' is given by the formula LTU/24,02 a,...

if the Munk airship formula is used for the point of action of the air, the following approximation is obtained:

Ja1=axial moment of inertia Jq=transverse moment of inertia e=angular velocity cm=coefcient of moments aboutthe tranverse axis of the projectile p=air density v=velocity of the projectile Fg=`crosssection of the projectile D=diameter of the projectile VG=volume of the projectile Ig=length of the projectile xs=distance of centre of gravity from the tip of the projectile It is an object of the present invention to utilize the accuracy of spin-stabilized projectiles Without having to accept at the same time the undesired effects on the operative hollow-charge portion or on some other spin-sensitive useful load, for example of ya supply projectile, of high rotational speeds. The invention provides means for limiting the speed of rotation as a function of the permissible peripheral speed of the hollow-charge or of some other spin-sensitive useful load and for keeping all 3,i44,828 Patented Aug. 18, 1964 ICC other quantities having an iniluence on the stability number of the projectile correspondingly large or small.

According to the invention, that is achieved by arranging the combustion chamber in annular form around the spin-sensitive, operative portion of the projectile andfixing the chamber to the portion in such mannernthat the centres of gravity of both chamber and portion coincide or are at only a short distance from one another, and that the resultant ratio of the moments of inertia Ja/Jq (axial moment of inertia/ transverse moment of inertia) is about 0.4 to 1.5.

Realization of a ratio of the moments of inertia Ja/ Jq of such a value also provides a very compact form of projectile, the spin-sensitive operative portion being 1ocated in the interior of the projectile and being surrounded by a combustion chamber of annular form.

As a further subsidia1y of the invention, a momentof-inertia ring, preferably consisting of a material having a high speciiic gravity, is at the centre of gravity of the projectile or at a small distance therefrom inside or outside the annular combustion chamber in order to increase further the ratio of the moments of inertia.

In conjunction with those constructional measures and in order to reduce the velocity of the projectile, which appears as a square in the denominator of the stability number equation given above, the thrust program of the rocket is so designed by employing a cruising propelling unit for compensating the drop in speed after the main propelling unit has burned out such that the projectile reaches a relatively low maximum speed. This can be achieved with advantage by arranging the cruising propelling unit adjacent the outer surface of the projectile where it exhibits the elect of the moment-of-inertia ring and can either replace or supplement that ring.

imbodiments of the invention will now be described in greater detail with reference to the accompanying drawings of which:

FlG. l is a section of a rocket projectile with a propelling unit,

FIG. 2 is a section of a rocket projectile with separate starting and cruising propelling units, and

FIG. 3 is a section of a rocket projectile with a starting propelling unit and a ram jet propelling unit as cruising propelling unit.

The rocket projectile illustrated in FIG. l comprises a central, spin-sensitive operative portion arranged with its axis along the longitudinal axis and shown as a hollow charge l. A combustion chamber 2 of annular form with a propellent charge 3 is arranged round the hollow charge l. Reference numeral idesignates nozzles which are also arranged in the form of a ring round the hollow charge at the rear end of the combustion chamber and are slightly inclined or skewed as shown to produce the spin. The structure of the projectile is such that the centres of gravity of the propelling unit and the hollow charge are substantially coincident and form a common centre of gravity S.

Within the combustion chamber there is arranged a displacement of centre of gravity and a change in the axial moment of inertia owing to the burning away of the propelling units and through the influence of the rocket velocity, a stabilityl number a 1 is obtained over the entire operating range.

It is furthermore possible to position the propellent charge of the cruising propelling unit in the common combustion chamber on the outer shell of the combustion chamber, so that, possibly with the use of ballast powder, it has the properties of the moment-of-inertia ring described and can increase the etIect thereof in the critical range of speed of rotation in order to maintain an adequate stability number, or can replace the moment-ofinertia ring.

A rocket projectile illustrated in FIG. 2 likewise has an operative portion consisting of a hollow charge 1. The operative portion is surrounded by a starting propelling unit 3 and a cruising propelling unit 6 disposed concentrically with respect to the hollow charge 1. The projectile shown in FIG. 2 is otherwise similar to that shown in FIG. 1, the positions of the centres of gravity of the hollow charge, the starting unit 3 and the cruising unit 6 being substantially coincident.

In the construction shown in FIG. 3, a cruising propelling unit 6 is arranged concentrically around the starting propelling unit 3 and is in the form a ram-jet propelling unit. The construction of the projectile of FIG. 3 is otherwise similar to that of the projectile shown in FIG. 1.

We claim:

1. A spin-stabilized rocket projectile comprising in combination a spin-sensitive portion, a rst annular combustion chamber forming a starting propulsion unit for said projectile, said irst chamber surrounding said portion and xed thereto, and a second annular combustion chamber forming a cruising propulsion unit for said projectile and also surrounding said portion, the centres of gravity of said portion and said first and second combustion chambers being substantially coincident and the ratio of axial moment of inertia to transverse moment of inertia of said projectile lying within the range of about 0.4 to 1.5.

2. A spin-stabilized rocket projectile comprising in combination a spin-sensitive portion, a rst annular combustion chamber forming a starting propulsion unit for said projectile, said rst chamber surrounding said por tion and fixed thereto, and a second annular combustion chamber forming a cruising propulsion unit for said pro- 4f jectile and positioned around said portion to act as a moment-of-inertia ring, the centres of gravity of said portion and said first and second combustion chambers being substantially coincident and the ratio of axial moment of inertia to transverse moment of inertia of said projectile lying Within the range of about 0.4 to 1.5.

3. A spin-stabilized rocket projectile comprising in combination a spin-sensitive portion, a rst annular combustion chamber forming a starting propulsion unit for said projectile, said lirst chamber surrounding said portion and xed thereto, a second annular combustion chamber forming a cruising propulsion unit for said projectile and also surrounding said portion, the centres of gravity of said portion and said first and second chambers being substantially coincident and the ratio of arial moment of inertia to transverse moment of inertia of said projectile lying Within the range of about 0.4 to 1.5, and a moment-of-inertia ring disposed around and concentric With respect to said portion in a position such that the centre of gravity of said ring is substantially coincident with said centres of gravity of said portion and said rst and second combustion chambers.

4.'A spin-stabilized rocket projectile comprising in combination a spin-sensitive portion, a starting propulsion unit including an annular combustion chamber surrounding said portion and fixed thereto, and a ram jet cruising propulsion unit including an annular combustion chamber surrounding the combustion chamber of said starting propulsion unit, the centres of gravity of said portion and the combustion chambers of said starting and said cruising propulsion units being substantially coincident and the ratio of axial moment of inertia to transverse moment of inertia of said projectile lying within the range of about 0.4 to 1.5.

References @item1 in the tile of this patent UNITED STATES PATENTS 2,519,878 Bjork et al. Aug. 22, 1950 FOREIGN PATENTS 5,099 Great Britain Dec. 12, 1878 516,818 Great Britain J an. l2, 1940

Claims

1. A SPIN-STABILIZED ROCKET PROJECTILE COMPRISING IN COMBINATION A SPIN-SENSITIVE PORTION, A FIRST ANNULAR COMBUSTION CHAMBER FORMING A STARTING PROPULSION UNIT FOR SAID PROJECTILE, SAID FIRST CHAMBER SURROUNDING SAID PORTION AND FIXED THERETO, AND A SECOND ANNULAR COMBUSTION CHAMBER FORMING A CRUISING PROPULSION UNIT FOR SAID PROJECTILE AND ALSO SURROUNDING SAID PORTION, THE CENTRES OF GRAVITY OF SAID PORTION AND SAID FIRST AND SECOND COMBUSTION CHAMBERS SUBSTANTIALLY COINCIDENT AND THE RATIO OF AXIAL MOMENT OF INERTIA TO TRANSVERSE MOMENT OF IN-