US3752425A

US3752425A - Self-propelled non-guided missiles

Info

Publication number: US3752425A
Application number: US00193093A
Authority: US
Inventors: B Detalle
Original assignee: Societe Europeenne de Propulsion SEP SA
Current assignee: Societe Europeenne de Propulsion SEP SA
Priority date: 1970-10-28
Filing date: 1971-10-27
Publication date: 1973-08-14
Anticipated expiration: 1990-08-14

Abstract

This invention relates to self-propelled non-guided missiles, of the type comprising a collapsible tail unit arranged on the external wall of the missile body close to the ejection nozzle for the propellant gases: a gyroscope is mounted in the missile body and has a shaft disposed in a plane perpendicular to the longitudinal axis of the missile, and a chamber to contain a suitable propellant fuel. In such a missile, the invention provides a lifting aerofoil that is secured to the external wall of the body and close to the plane containing the centre of gravity thereof; the gyroscope acts as a banking stabiliser to maintain a substantially constant attitude for the aerofoil and is mounted in a body which is locked by means of a locking member until the propellant fuel has been combusted.

Description

United States Patent 11. 1

ll'ietaile SELF-PROIFELLIED MGM-GUIDED Mllll11 [75] inventor: Bernard Andre llilenri llletaile,

Llillaye les-lltoses, France 22 Filed: on. 27, 19711 21 App]. No: 11%,1193

[30] Foreign Application Priority Data 1451 Aug. 11-11, 11117:;

FOREIGN PATENTS OR APPLICATKONS 866,598 8/ 1941 France 244/328 Primary Examiner-Benjamin A. Borchelt Assistant Examiner-James M. Henley A ttorney- Lorimer P. Brooks, G. Thomas Delahunty et a1.

[ 7 ABSTRACT This invention relates to self-propelled non-guided missiles, of the type comprising a collapsible tail unit an ranged on the external wall of the missile body close to the ejection nozzle for the propellant gases: a gyroscope is mounted in the missile body and has a shah disposed in a plane perpendicular to the longitudinal axis of the missile, and a chamber to contain a suitah1e propellant fuel. In such a missile, the invention provides a lifting aerofoil that is secured to the external wall of the body and close to the plane containing the centre of gravity thereof; the gyroscope acts as a banking stabiliser to maintain a substantially constant attitude for the aerofoil and is mounted in a body which is locked by means of a locking member until the propellant fuel has been combusted.

6 Claims, 5 Drawing Figures PATENTEU MIC 14 I975 SHEET 1 0F 2 mw mm s. R

SHEET 2 0F 2 PAIENIEDAus I4 nun I SELF-PROPELLED. NON-GUIDED MISSILIESv The present invention. relates to a self-propelled missile of the rocket-type.

Unarmoured non-guided rockets are generally used to reach targets located at one or two kilometres from the point of launching. The trajectory-of the rockets is ballistic, i.e-., having the shape of a parabola whose concavity is uppermost. For this reason, the trajectory has a highest point which it isnecessary to take into account in order to have good firing accuracy. In order to improve the latter feature, the missile is propelled so that, at the outlet or muzzle of the launching tube, it has an initial velocity as great as possible whereby its average speed over the trajectory is improved. On the other hand, attempts have been made to increase the firing accuracy by increasing the missiles velocity, thus decreasing its flight time; as a result, there is a decrease in the highest point of the trajectory. Nevertheless, with the stabilisation and propulsion devices hitherto fitted to, known rockets, a given highest point has not been able to be maintained beyonda certain range.

This has led to the production of propellants,- either which are at the limit of technological possibilities, signifying less viable constructions which are very expen-- sive, or to stabilising the said missiles, for example by a collapsible tail unit associated possibly with a pitch stabilising gyroscope. Nevertheless, these stabilising means have no effect on the highest point of trajectory of the missile, the centre of gravity of which describes a ballistic trajectory.

The gyroscope may be used in the form of an inertia platform insofar as it is an attitude reference, the movement of the gyroscope frames acting only on steering means in order to bring the missile into a predetermined trajectory.

The. present invention has for an object the removal or reduction of the aforementioned drawbacks and proposes an improved non-guided rocket which has a rectilinear trajectory in the useful part of its application and which keeps a substantially constant banking attitude on its trajectory. I

To this end, the invention consists in a missile of the type comprising a collapsible tail unit arranged on the external wall of the missile body close to the ejection nozzle of the gases, and having a gyroscope mounted in the missile body and has a shaft disposed in a plane perpendicular to the longitudinal axis of the missile, wherein a lifting aerofoil is fixed to the external wall of the body of the missile and close to the plane containing the centre of gravity of the said missile, and the gyroscop'e acting as a banking stabiliser tomaintain a substantially constant attitude to the said lifting aerofoil and being'mounte d in a body which is locked by means ofa locking member until the end of combustion of the propellant.

One advantage of the present invention resides'in the fact that the rocket reaches the sighted target whilst maintaining a rectilinear trajectory and in the fact that the gyroscope is used to stabilise the missile when banking so as to oppose any disturbance which tends to cause the rocket to swing about its axis of banking, an inertia such that the starting attitude is not modified.

According to another feature of the invention, the lifting aerofoil is collapsible and, in the collapsed state, is located within the launching tube and is mounted on a support member which is hinged at one of its ends on a shaft secured to or integral with the body of the missile and. the other end of which is capable of moving in a vertical plane with respect to the body of the said missile; moreover, a compensating spring is interposed between the support member and the body of the missile.

As a consequence, the rocket can be used in a wide range of temperatures. In fact, taking into account its possible application, the rocket may be launched in regions where the temperatures are extremely different. Thus, the rocket can be launched in a region which is at a temperature of the order of 50 C or in another region the temperature of which may be equal to C or 50 C. However, it goes without saying that the specific weight of the air in such different regions varies. There thus follows therefrom a variation in the lift which was provided to balance the weight under normal conditions of use. In fact, it is known that the lift P of a missile or of a craft is proportional, on the one hand, to the specific mass p of air and to the angle of incidence i which an aerofoil makes with the body of the missile. The present invention thus enables this angle of incidence i to be varied to compensate for a possible change in this specific mass p of the air and means that the lift P is always equal to the weight whatever the ambient atmospheric conditions.

According to another feature, the frame of the gyroscope is locked until the end of combustion of the accelerating propellant. This avoids the gases employed for driving the rotor of the gyroscope causing deviation of the frame, which deviation would cause a loss in speed of the gyroscope and would not allow the latter to assume a position appropriate for maximum stabilisation in banking.

Other advantages and features of the invention will become apparent from a study of the ensuing description taken in conjunction with the accompanying drawings'which show an embodiment thereof purely by way of non-limiting example, and in which:

FIG. 1 is a view in longitudinal partial section of one embodiment of the invention,

FIG. 2 is a view in section along the line 11-11 of FIG.

FIG. 3 is a view in section along the line IIi-lll of FIG. 2,

FIG. 4 is a view in perspective of the lifting aerofoil mounted on its support, and

FIG. 5 represents half-sections along the lines V-V and Vl-Vl of FIG. 2.

Referring now to the drawings, there is shown a rocket 1 which comprises two stages 3 and 4 for launching or acceleration, and flight respectively, intended to preserve, during the trajectory of the rocket, the velocity with which it was fired at the nozzle of the launching tube. In the acceleration stage 3 is arranged a propellant 5, for example a powder cake, capable of producing by its combustion in the selected embodiment, a thrust of the order of 40.000 N for approximately 6 X 10 seconds. The powder cake 5 is ignited by an ignitor 7 arranged between the stages 3 and 4. In the flight stage 4 is arranged a powder cake 6 designed to ensure a thrust of the order of 400 N whilst the suitable military load is arranged in the head 2 of the rocket. The rocket 7 ignites the powder cakes 5 and 6 simultaneously and is formed by a propellant load arranged at the rear of a chamber 8 extended by a conduit 9 ending at its open end 9a in an ejection nozzle 9b intended for the evacuation of the combustion gases from the flight stage. In the wall of the chamber 8 are arranged calibrated orifices 11 which enable the ignition of the powder cake 6 of the flight stage 4 to be produced. An ejection nozzle 12 is provided in the usual manner at the rear of the rocket 1.

The rocket 1 also comprises a collapsible tail unit 13 formed for example by four stabilising blades 14 mounted at 90 from one another on the external wall of the nozzle 12. Each stabilising blade 14 is fixed to a shaft 15 by means of fixing lugs 16, the said shaft 15 being arranged in a socket 17 mounted in recesses arranged in projecting appendices 18 provided on the external wall of the nozzle 12.

According to one feature of the invention, the rocket 1 is fitted with a lifting aerofoil 19 shown in detail in FIG. 4, situated close to the plane containing the centre of gravity of the rocket and formed by two wing stubs or'ailerons 20 each mounted on a shaft 21 carried in flanges 22 arranged on a support 23, the wing stubs 20' being locked in theopen position.

The support 23 is a dished member having in crosssection the shape of an inverted U and the arms of which serve as support flanges for two

shafts

24 and 26 mounted respectively on a front cover 25 and arear cover 28, both integral with or secured to the body of the missile (FIG. 5). A guide slot 27 is arranged in the rear cover 28, through which the shaft 26 passes, whereby the support 23 and consequently the lifting aerofoil 19 are capable of moving simultaneously in a vertical plane with respect to the plane of the body of the missile 1 about the shaft 24 and in such a manner as to cause the angle of incidence i to vary. In fact, as mentioned above, the lift P of the rocket which balances its weight is a function of the specific mass p of the air and of the angle of incidence i which the lifting aerofoil 19 makes with the body of the rocket 1. Thus, if in a region of use p increases,. the support 23 is caused to pivot about the shaft 24 so as to decrease the angle of incidence i in order to preserve a constant value for the lift 1.

It goes without saying that the guide slot 27 may be arranged on the support 23 as is shown at 27a on FIG. 4, the shaft 26 thus being fixed and mounted on the rear cover 28.

A compensating spring 29 fixed by any suitable means such as rivets 30 is placed between the inner face 23a of the support 23 and the body of the rocket '1 (FIG. 1). This compensating spring 29 is calibrated so that it tends to return the lifting aerofoil 19 to an incidence i such that the weight of the rocket is always balanced by the lift. A resilient washer 31 for damping vibrations may also be interposed between the compensating spring 29 and the face 23a of the support 23.

In an enclosure 32 arranged in the body of the rocket 1 and in front of the flight stage 4 is arranged a banking stabilisation gyroscope 33, the shaft 34 of which is situated, on the one hand, in a plane perpendicular to'the longitudinal axis of the rocket and, on the other hand, in the vertical median plane of the lifting aerofoil. The shaft 34 of the gyroscope is mounted in bearings 35 integral with or secured to a body 36 fixed by bearings 37 on the inner wall of the body of the rocket. The rotor of the gyroscope is formed by a vaned wheel 38 (the vanes are not shown) and which is driven by expanded gases issuing from a gas collector 39 which is arranged in a sector in the enclosure 32 either in the form of a conduit arranged in the horizontal plane containing the vaned wheel of the gyroscope or in the form ofa spherical cover as is shown on FIGS. 1 and 2 and which comprises awall 40 pierced with distribution orifices 41. The gas collector 39 is connected to at least one annular input chamber 42 when the gas collector 39 is in the form of a spherical cover, by means of at least one expansion nozzle 43. The annular input chamber 42 is connected by a supply conduit 44 to the combustion chamber of the acceleration propellant 5.

The body 36 of the gyroscope 33 is locked (FIG. 5) until the end of combustion of the acceleration propellant 5 by means of a pyrotechnical delay charge 45 which is located in a bore 46 ofa cylinder 47 and which acts on a piston 48 housed partly in the bore 46 and in an orifice 49 made in a cap 50 of the body 36.

I claim:

1. A self-propelled non-guided missile of the type comprising a collapsible tail unit arranged on the external wall of the missile body close to the ejection nozzle for the propellant gases, a gyroscope mounted in the missile body so as to have the shaft of its rotor arranged in a plane perpendicular to the longitudinal axis of the missile and its rotor driven by a part of said propellant gases, at least one propellant charge stage located in said missile body and communicating with said ejection nozzle, and also comprising close to the gravity center of the missile body a lifting aerofoil secured to the external wall of the missile body, said lifting aerofoil being mounted on a support member which by one of its ends is hinged to a support member shaft, the support member shaft being perpendicular to the plane defined by said longitudinal axis of the missile body and the shaft of the gyroscope rotor and secured to said missile body, and a compensating spring located between and secured to said support member and to said missile body so as to limit the angle of incidence which said lifting aerofoil makes with said missile body.

2. A missile according to claim 1, wherein said lifting aerofoil is collapsible.

3. A missile according to claim 1, wherein said support member comprises two flanges situated in different planes parallel to said plane comprising said longitudinal axis of said missile, said flanges being locked together by a shaft perpendicular to said flanges and arranged for vertical movement and mounted in the guide slot of a cover secured to said missile body.

4. A missile according to claim 1, wherein the frame of the gyroscope is hinged in aplane perpendicular to the axis of rotation of said gyroscope rotor on bearings secured to the missile body and the frame is locked by means of a delay pyrotechnical charge arranged to be fired simultaneously with said propellant charge.

5. A missile according to claim 1, wherein the support member comprises two flanges situated in different planes parallel to the plane comprising the longitudinal axis of the missile body and the shaft of the gyroscope rotor, each of said flanges being slotted to receive a common shaft perpendicular to said flanges and mounted in a cover secured to the missile body.

6. A missile according to claim 1, wherein a resilient washer is interposed between the compensating spring and the support member.

*9 III k k

Claims

4. A missile according to claim 1, wherein the frame of the gyroscope is hinged in a plane perpendicular to the axis of rotation of said gyroscope rotor on bearings secured to the missile body and the frame is locked by means of a delay pyrotechnical charge arranged to be fired simultaneously with said propellant charge.