FR2572511A1 - ROCKET OR RACKET STABILIZER - Google Patents

Abstract

THE INVENTION CONCERNS A ROCKET STABILIZER INCLUDING A CIRCULAR (OR TUNNEL) FIN 6 AERODYNAMICALLY PROFILE SLIDING AXIALLY ON THE TUBE 1 OF THE ROCKET 2 SO AS TO REMOVE IT AT THE TIME OF LAUNCHING TO LEAVE PART OF L 'EXTERIOR AIR ENTER THE AILERON AND EXERCISE A DIRECTIVE OR COMPENSATING ACTION. THE DEVICE, BY ITS ZERO DIMENSIONS, ALLOWS LAUNCHING BY TUBE OF ROCKETS WITH MAXIMUM OUTLET DIAMETER.

Description

d-j? 2572511 The present invention relates to rocket stabilizers

  and especially, but not exclusively, a stabilizer for

tube launched rocket.

  The stability in flight of a rocket requires that its center of pressure

  sion is located behind its center of gravity, at a distance called "static margin" which, preferably, will not be less than half the diameter of the rocket body. This essential static margin is normally obtained by bringing

  the center of pressure at the desired point by means of surfaces

  lanyards or stabilizers attached to the rear of the rocket. The use of drag stabilizers comprising a flared rear part or stabilizing control surfaces for subsonic flight is known, but these devices, as their name suggests, create a drag and are not suitable for flight

  transonic or supersonic because the directional errors

  possible are too large. Door stabilizers

  tance, comprising fixed axial empennages projecting radially

  of the rocket's propellant nozzle, are known to be suitable for all flight speeds but, for rockets launched by tube of caliber corresponding to the rocket body,

  the projection of the tail will necessarily reduce the mini section

nozzle outlet male.

  The shrinkage thus caused can be reduced by use

  folded tail that deploys only after launch.

  A known form of folded empennage is that of the empennages known as "penknife blades" in which several fins are pivotally mounted on one end of the nozzle and folded forward along the latter while it is

  in the launch tube. These empennages are necessarily

  narrow, have a high aerodynamic elongation and cannot

  wind not provide effective stabilization at higher speeds

  lower than subsonic velocities.

  A form of empennage better suited to trans-

  sonic is that of the circumferentially folded tail-

  lying and abruptly opening out of the launch tube

  to form fins of low elongation, but these em-

  5.pennages also require a certain space, which necessarily reduces the outlet section of the nozzle. In addition, these empennages and their spring release mechanism are quite complex and heavy and cause an unwelcome shift of the center of gravity of the rocket towards the rear. Moreover, the

  various fins constituting the empennage must be synchro-

  to open simultaneously.

  Another form of fixed lift stabilizer is that

  affecting the shape of a cylinder mounted coaxially on the outside

  back side of a bomb, for example, but, as in the &

  previous examples, the diameter of the part carrying the sta-

  stabilizer should be less than the rest of the body of the

  projectile we are considering launching by tube.

  The object of the present invention is to provide a stabilizer

  effective rocket at all trajectory speeds,

  dragging that little or no reduction in the cross-section of the pipe

  era, and thus allowing to equip with a heating nozzle

  bre essentially maximum a rocket launched by tube.

  The rocket stabilizer according to the invention, for a rocket whose nozzle has an axially symmetrical outlet flange, comprises an aerodynamically profiled annular part or circular fin arranged so as to be aligned, during the

  trajectory, on said outlet edge and at a certain distance

  axial flow downstream of it so as to define with it a circumferential opening, the circular fin having a

  leading edge opposite the outlet rim and an insulating surface

  dull offering a longer flow path than its sur-

external face.

  Preferably, the external part of the nozzle will be shaped,

  in the vicinity of the outlet rim, so as to go into sta-

  does it scale slightly inward downstream so

  assist, during the flight of the rocket, in the intake of air ex-

  térieur to the internal surface of the circular fin.

  The circular fin can be fixed on the nozzle at a certain

  taine axial distance from the latter by means of a chassis

  tor with the appropriate openings.

  But in a preferred arrangement, in which the circular fin does not deploy until after launch, the supporting chassis is arranged so as to slide axially on the rocket barrel so as to allow the leading edge of the fin to be closed by the outlet edge. In such an arrangement, the outlet rim and said leading edge are shaped so as to fit closely together, and the rest of the internal surface of the circular fin is shaped so as to continuously extend the internal contour, it ie the outlet cone, of the nozzle. Before launching, the circular fin is closed by the flange / outlet, which excludes any discontinuity in the outlet cone which could

  cause deviations from management during launch.

  During the combustion period and while the nozzle is still in the launch tube, the east circular fin

  held forward against the exit lip by pressing

  exhaust, but after leaving the tube, the outside air flow causes the circular fin to deploy and its

  carrier frame rearward in active position.

  The invention is described below in detail with reference to a preferred, non-limiting example of embodiment shown in the accompanying drawings in which:

  - Figure 1 is an axial section of a fuel stabilizer

  according to the invention, comprising a circular fin

  foldable shown in the position before the launcher

  is lying; - Figure 2 is a partial axial section of the stabilizer

  in Figure 1 showing the circular fin deployed in po-

  flight location; and - Figure 3 is a cross section of the stabilizer along line A-A of Figure 1.

  The rocket stabilizer shown in Figures 1 and 3 includes

  takes a propellant nozzle 1 fi3e to the body 2 of a rocket and

  comprising a narrowing 3, an outlet cone 4 and a. re-

  exit edge 5. A light annular aerofoil piece

  dynamically (or circular or barrel fin) 6, made of aluminum alloy for example, comprising a leading edge 7, an external surface 8 and an internal surface 9, is carried, downstream of the rim 5, on the ends of eight rods 10 parallel to the axis of the rocket, regularly distributed around the circumference of said rim - the other ends of the

  luit rods being fixed to a stop ring 11 mounted slidingly

  health on the rocket body 2. The rods 10 can, in their middle part, slide through a bead 12 fixed to the propellant nozzle 1 and able to penetrate, in adjustment

  sliding, in the launch tube (not shown).

  The outer surface of the nozzle 1 has a rear part

  1st 13 slimming inwards and extending

  dant downstream of the outlet edge 5, itself shaped so as to match a part of the leading edge 7 of the circular fin 6 when the craft is in the position ready for launch. The inner surface 9 of the circular fin 6 is shaped so as to be an extension of the last supper

output 4.

  When the device is deployed (see Figure 2), the stop ring 11 slides downstream to come to rest against

  the bead 12 against which it remains held by the

  flow of outside air and fitting of a calandering ring

  te (or wedge-shaped profile) 14 fixed on the upstream face

  of the bead with an annular opening 15 of corresponding profile

'2572511

  laying in the downstream face of the stop ring 11.

  The circular fin 6 therefore moves away from the nozzle 1 towards

  downstream to discover a circular opening 16.

  During the flight of the rocket, the outside air flowing along the body 2 and the nozzle 1 is divided at the location of the outlet flange 5, a part of this air continuing its movement along the outside surface 8 of the circular fin while

  that the rest of the air enters through opening 16 to flow

  along the inner surface 9 of said circulating fin

  re, which generates a resulting acting radial pressure,

  through all the points of the annular pressure center 17 of the

  circular leron 6, in the direction of the axis 18 of the rocket o The center of pressure of the entire missile is therefore displaced towards the rear by the action of the circular fin 6 during flight. Increasing the static margin accessible this way

  does not only depend on the wing area of the circling fin

  ring 6 but also the proportion of the air flow a-

  led to flow along its inner surface 9. This é-

  internal air flow itself depends on the axial length of the opening 16, the external taper angle of the nozzle

  13 and the shape of the leading edge 7 of the circular fin.

  All these variables are chosen so as to remain within the

  limits of dimensional easements imposed by condi-

  launches, in order to obtain an optimal distribution

flows.

  It is obviously important to maintain a precise axial symmetry

  of all aerodynamically profiled components, all er-

  alignment error which may result in a reduction of the angle

  regulation and therefore generate di-

  rection unless the rocket equipped with this stabilizer

  be subjected to a slow gyratory movement.

  It will appear to specialists in the field that the present

  invention admits various variants. The circular fin de-

  foldable, for example, can be carried by a cylindrical sleeve

  slide sliding on the rocket body and pierced with a crown-

  no rectangular orifices giving after deployment the opening

  circumferential size necessary but strong enough to immo-

  reliably stabilize the circular fin. In another embodiment, this sleeve may be wedged in the deployed position if it turns out to be of no importance for the launching conditions which the circular tailgate occupies, from before departure,

its deployed position.

  It will also be possible to improve the air flow on the internal surface of the circular fin by providing a

  small profiled ring, sort of forming a compensating flap

  sater, bent outward, directed upstream and fixed,

  by its downstream end, at the leading edge of the circulating fin

  air, so as to be held against the nozzle while it

  is in the launch tube but protrudes out

  both of the tube so as to form a small conical flow guide

lement.

  The deployment system is also likely to vary

  your. For example, a circular fin with a chassis

  suitable tor can be placed separately in the launch tube

  cement upstream of the rocket so that at the time of departure

  the latter passes through the circular fin and the ac-

  crooked by its nozzle bead before leaving. This provision

  sitive may have advantages with a telescopic launch tube. All these devices make it possible to produce a light weight and space-saving rocket stabilizer, including the

  deployment has no springs and locks automatically

mattress.

Claims (11)

  1. Rocket stabilizer, for a rocket whose nozzle has an axially symmetrical outlet flange, characterized by   what it includes an aerodynamically pro-   spun or "circular fin" arranged so as to be   hinged, during the flight of the rocket, on said exit flange and at a certain axial distance downstream from it so as to   define with it a circumferential opening, said ai-   circular protrusion having a leading edge opposite said   outlet rim and an interior surface providing   flow course longer than its outer surface.   2. Rocket stabilizer according to Claim 1, characterized   in that the nozzle and the circular fin have essential-   the same outside diameter, and in that the nozzle has   sedes an outwardly tapering portion tapering in the vicinity of the outlet rim and reducing, at the location of said outlet rim, to an outside diameter less than   that of the leading edge of the circular fin.   3. Rocket stabilizer according to any of the Revendica-   1 or 2, characterized in that the circular fin is   mounted on an axially symmetrical supporting frame.   4. Rocket stabilizer according to Claim 3, characterized in that the supporting frame is rigidly fixed to the nozzle   5. Rocket stabilizer according to any of the Revendica-   tions 3 or 4, characterized in that the supporting frame comprises   carries a coaxial cylinder extending upstream from the   leading edge of the circular fin and provided with a of radial openings.   6. Rocket stabilizer according to Claim 3, character-   in that the supporting frame is slidably fixed   te to the nozzle to allow the axial closure of the leading edge of the circular fin by the outlet flange when the rocket is not yet launched.   7. Rocket stabilizer according to Claim 6, character-   se en oe that the nozzle is provided with a bead pierced with a plurality of axially parallel holes, and in that the carrier frame comprises a plurality of axially parallel rods, diac rod can slide in one of the holes   and being fixed by one of its ends to the circular fin   air and by its other end to a stop ring for before sliding on the rocket.   8. Rocket stabilizer according to any of the Revendica-   1, 2, 3, 4, 5, 6, 7 or 8, characterized in that the inner surface of the circular fin is shaped so as to continuously extend the inner contour of latuyere.   9. Rocket stabilizer according to any of the Revendica-   1, 2, 3, 6, 7 or 8, characterized in that the protruding edge is shaped so as to push the edge exactly of attack on the circular fin.   10. Rocket stabilizer according to Claim 3, character-   in that before launching the circular fin and the supporting frame are arranged axi-symmetrically in a   launch tube upstream of the nozzle, said chassis   tor being captured, during launch, by a bead carried by the nozzle.   11. Rocket stabilizer according to any of the Revendica-   1, 2, 3, 4, 5, 6-, 7, 8 or 9, characterized in that a flexible conical guide facing upstream is fixed   at the periphery of the leading edge of the circular fin.