Note: Descriptions are shown in the official language in which they were submitted.
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This invention relates to rocket stabilisers and in
particular but not exclusively, to a stabiliser for tube-
launched rockets.
For a rocket to be stable in flight it is essential
that its centre of pressure should lie aft of its centre
of gravity by a distance, known as the static margin, of
preferably not less than one half of the rocket body
diameter. Normally, in order to achi.eve this necessary
static margin, the centre of pressure is brought to the
desired position by means of extended surfaces or
stabilisers fitted at the tail of the rocket.
: Drag stabilisers comprising a flared tail or
streamers for example, are known for subsonic flight but
these, as their name implies, impart drag and are not
suitable for transonic or supersonic flight as the
directional errors that could occur are too great.
Lifting stabilisers comprising fixed axial fins
projecting radially outwards from the rocket nozzle are
known to be suitable for all flight speeds but, if the
rocket is to be launched from a launch tube of calibre
matched to the rocket body, the required protrusion of
; the fins necessarily constricts the nozzle throat.
The degree of constriction engendered can be reduced
by the use of fold-away fins which are deployed only
after launching. One known configuration of fold-away
fins is that of penknife fins, a plurality of which are
pivoted at one end to the nozzle and folded forward
alongside the nozzle when in the launch tube. Such fins
are essentially narrow, have a high aspect ratio and
cannot provide effective stabilisation at flight speeds
above subsonic. Another known configuration more
suitable for transonic flight is that of wrap-around fins
which spring open on emergence from a launch tube to
provide fins of lower aspect ratio, but even these fins
require some stowage space which necessarily constricts
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the nozzle throat. Furthermore such fins and their
spring release mechanisms are comparatively complex and
also heavy, causing the centre of gravity of the rocket
to be shifted disadvantageously rearwards. In addition
the fin blades may need to be synchronised for
simultaneous opening.
One other configuration of a fixed lifting
stabiliser is that known as a "ring-tail", i.e. a
cylinder mounted coaxially at the downstream end of a
bomb, for example, but as in the previous examples, the
portion supporting the stabiliser must be of smaller
diameter than the rest of the body if tube-launching is
- intended.
The present invention seeks to provide a rocket
stabiliser effective at all flight speeds which will
engender little or no constriction of the nozzle throat
and consequently permit a substantially full-calibre
nozzle to be employed with a tube-launched rocket.
According to the present invention, a rocket
stabiliser for a rocket having a nozzle with an axi-
symmetric exit lip includes an annular aerofoil arranged
; in flight to be axially aligned with and axially spaced
downstream from the exit lip so as to define therewith a
circumferential aperture, the annular aerofoil having a
leading edge opposed to the exit lip and an inner surface
of greater flowpath length than that of its outer
surface.
Preferably the rocket nozzle is externally contoured
; adjacent the exit lip so as to taper slightly inwardly in
the downstream direction in order to assist in-flight
admission of exterior airflow to the inner surface of the
aerofoil.
The aerofoil may be rigidly fixed to the nozzle in
axially spaced relationship thereto by means of a
suitably apertured support frame.
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In a preferred arrangement however, in which the
aerofoil is not deployed until after launch, the support
frame is arranged to be axially slideable upon the rocket
nozzle so as to permit the leading edge of the aerofoil
to be closed with the exit lip. In such arrangement the
exit lip and the leading edge are shaped to mate
precisely and the remaining inner surface of the aerofoil
is shaped to form a smooth continuation of the inner
contour, i.e. exit cone, of the nozzle. Prior to launch,
the aerofoil is closed with the exit lip, thus ensuring
that no discontinuities are present in the exit cone
which might impart unwanted directional deflections upon
firing. During the rocket burn time and whilst the
~ocket nozzle is still within the launch tube, the
aerofoil is held forward against the exit lip by the
rocket exhaust pressure but, upon leaving the tube, the
external airflow acts to deploy the aerofoil and its
support frame rearwards to the operative position.
An embodiment of the invention will now be described
~ 20 by way of example only, with reference to the
accompanying drawings of which
Figure 1 is an axially sectioned view of a rocket
stabiliser having a deployable annular aerofoil
illustrated in the pre-launch position,
Figure 2 is an axially sectioned part-view of the
same rocket stabiliser with the aerofoil deployed in the
flight position, and
Figure 3 is a radially sectioned view of the same
rocket stabiliser taken on the line A-A of Figure 1.
The rocket stabiliser illustrated in Figures 1 and 3
comprises a rocket nozzle 1 secured to a rocket body 2
and having a throat 3, exit cone 4, and exit lip 5. A
light-weight annular aerofoil 6, of aluminium alloy for
; example, having a leading edge 7, an outer surface 8 and
an inner surface 9, is supported downstream of the lip 5
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upon one end of each of eight axially parallel rods 10
evenly disposed around its circumference, the other ends
of which are secured to as top ring 11 slideably mounted
upon the rocket body 2, the rods being intermediately
slideable through a bourrelet 12 which is secured to the
- rocket nozzle 1 to provide a sliding fit in the launch
tube (not shown). The stop ring 11 is of smaller
external diameter than the bourrelet 12.
The outer surface of the nozzle 1 has an inwardly
tapering rear portion 13 extending downstream to the exit
lip 5 which lip is itself contoured to mate with a
portion of the leading edge 7 of the aerofoil 6 when in
the pre-launch position. The inner surface 9 of the
aerofoil 6 is contoured to form a continuation of the
exit cone 4.
Upon deployment (see Figure 2) the stop ring 11
slides downstream to rest against the bourrelet 12 where
it is held thereafter by external air flow and by the
engagement of an annular wedge 14 located on the upstream
face of the bourrelet, with a correpsonding annular
wedge-shaped aperture 15 located in the downstream face
of the stop ring 11. The aerofoil 6 is thus displaced
downstream from the nozzle 1 to provide a circumferential
aperture 16.
In flight, external air flowing along the length of
the body 2 and the nozzle 1 divides at the exit lip 5,
some continuing along the length of the outer surface 8
of the aerofoil and the remainder entering the aperture
16 to flow along the longer inner surface 9 of the
aerofoil thus engendering a resultant radial pressure
acting through all points of the annular centre of
pressure 17 of the aerofoil 6 in a direction towards the
rocket axis 18. The centre of pressure of the overall
rocXet assembly is thus moved rearwards by the in-flight
action of the aerofoil.
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The increase in static margin achievable by this
means is not only dependent upon the surface area of the
aerofoil 6 but also upon the proportion of airflow
induced to flow along its inner surface 9. This inner
airflow is, in turn, dependent upon ~he axial length of
the aperture 16, the external taper angle of the nozzle
; portion 13, and the shape of the leading edge 7 of the
aerofoil. All of these variables are selected, within
the dimensional constraints imposed by the launch
requirements, to optimise flow partition.
It is of course important to maintain precise axial
symmetry of all parts of the aerofoil assembly, as any
alignment errors will result in a small trim angle and
hence cause directional errors, unless the rocket to
which the stabiliser is applied is subjected to slow
spin.
It will be apparent to those skilled in the art that
various other arrangements of the present invention are
possible. For example, the deployable aerofoil may be
supported upon a cylindrical sleeve which is slideable
upon the rocket body, the sleeve being apertured with a
ring of rectangular windows to provide the necessary
: circumferential aperture when deployed combined with
sufficient strength to locate the aerofoil securely.
Alternatively, such sleeve may be rigidly fixed in the
deployed position if any interference with launch
conditions which may be engendered by such a pre-
positioned aerofoil is not critical.
An additional means for improving the airflow to the
inner surface of the aerofoil is the provision of a small
outwardly flexed, upstream facing, flap-type collar,
attached at its downstream edge to the leading edge of
the aerofoil, such that it will be held against the
nozzle whi].st in the launch tube but will protrude to
form a small conical flow guide upon emergence from the
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tube.
Variations in the method of deployment are also
possible. For example, an annular aerofoil having an
appropriate support frame may be separately located in
the launch tube upstream of the rocket so that, on
firing, the rocket will launch through the aerofoil,
capturing it at the nozzle bourrelet before departure.
Such arrangement may be conveniently adopted when the
launch tube is of a tslescopic variety.
All these arrangements are capable of providing a
light-weight rocket stabiliser requiring little stowage
space, which may be deployed without springs and is self-
locking.
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