Release apparatus for spin stabilized self-propelled projectiles

A release mechanism for facilitating launching a spin-stabilized self-propelled missile. A missile support includes a rotary turbine having a receptacle defining a spin axis, and a fixed support for supporting the rotary turbine for rotation about the spin axis and for movement axially of the spin axis. A nozzle assembly extends from the missile into the receptacle and includes a fusible joint for heating by high temperature exhaust gases expelled by the missile to release the missile, and an aft nozzle section movable away from the missile on fusing and separation of the fusible joint. An abutment on the rotary turbine is disposed in the path of movement of the aft nozzle section for striking by the aft nozzle section to effect rapid movement of the rotary turbine and receptacle axially away from the missile on fusing and separation of the fusible joint.

FIELD OF THE INVENTION 
This invention generally relates to a projectile release mechanism and, 
particularly, to a release mechanism for facilitating launching a 
spin-stabilized self-propelled missile. 
BACKGROUND OF THE INVENTION 
It has become increasingly important to eliminate the features associated 
with a ballistic trajectory ordinarily followed by rockets and other 
jet-propelled projectiles, by forming the projectiles as spherical 
spin-stabilized missiles. The spherical missile spins about an axis 
upwardly inclined relative to the intended straight line path of flight 
and aligned with the missile propulsion thrust axis. The missile is 
released following ignition or activation of the propulsion system within 
the missile. The propulsion is effected by the reaction of the exhaust jet 
of, for example, a rocket motor housed within the spherical missile shell. 
Often such spherical spin-stabilized missiles are provided in conjunction 
with attachments secured to the front end of an assault weapon such as a 
rifle. 
Such spin-stabilized spherical self-propelled missiles experience 
difficulties in remaining stabilized during attainment of desired 
rotational speed and in coordinating the spinning and release of the 
missile. Release of the missile prior to attainment of adequate rotational 
speed can result in unstable flight. Delay of release after attainment of 
adequate rotational speed can result in a loss of propulsive range. 
Consequently, attempts have been made to provide means for temporarily 
restraining and automatically releasing a spin-stabilized self-propelled 
spherical missile during spinup. For instance, in U.S. Pat. No. 3,245,350 
to J. A. Kelly, dated Apr. 12, 1966, a mechanical release is provided 
between a rifle barrel and a spin-stabilized spherical missile in order to 
selectively release the missile. However, precise automatic release is not 
afforded. More specifically, U.S. Pat. No. 3,554,078 to Joseph S. Horvath, 
dated Jan. 12, 1971, provides a fusible link for temporarily restraining 
and automatically releasing a spherical spin-stabilized missile during 
spinup. Release of the spherical rocket missile from its rotary supporting 
means is effected by causing hot missile rocket exhaust gas to weaken by 
heating or to heat and soften or melt a separate fusible link member 
which, prior to weakening by softening or melting, secures the missile to 
the rotary support means. In this patent, the separate fusible link member 
is of the nature of a brazing alloy serving as one part of a nozzle 
assembly to secure the rocket to the rotary support means. The fusible 
link member is brazed between two separate fore and aft nozzle portions 
which are permanently secured to the missile and to the support means, 
respectively, as by threaded engagements. 
An improvement on the aforementioned prior art is disclosed in U.S. Pat. 
No. 4,395,836 to Baker et al, dated Aug. 2, 1983 and assigned to the 
assignee of this invention, wherein a new and improved nozzle assembly is 
disclosed. The nozzle assembly includes a unitary nozzle member having 
fusible joint means formed integrally therewith, between the missile and 
the rotary support means, thereby eliminating the assembly and brazing 
operations of prior devices as shown in the Horvath patent, and thereby 
considerably reducing manufacturing costs and improving accuracy. However, 
in this patent the fore and aft sections of the unitary nozzle, forwardly 
and rearwardly of the fusible joint means, are permanently fixed to the 
missile and to the support means, respectively, as by threaded 
engagements. 
Further improvements are shown in U.S. Pat. No. 4,403,435 to Baker et al, 
dated Sept. 13, 1983 and assigned to the assignee of this invention, 
wherein a further new and improved nozzle assembly includes projectile 
support means having open-ended receptacle means out of which fore and aft 
sections of the nozzle can move on fusing and separation of the fusible 
joint means. This patent also shows an improved register section for the 
missile or nozzle which is generally conical in configuration to improve 
alignment of the missile with the spin axis during initial separation of 
the fusible joint means. 
The present invention represents somewhat of a radical departure from the 
prior art in that a mass is caused to be urged or propelled rearwardly by 
the gases of the missile or separate or combined other force generating 
mechanism to strike an abutment means on the turbine or rotary means for 
the missile to cause the rotary means in its receptacle, to move rapidly 
away from the missile after separation of the fusible joint means. The 
present invention thus allows positive missile retention by the launch 
system rotary means during coupling fusing and therefore eliminates 
pointing error tip off forces initiated by the coupling fusing in any of 
the prior art. 
SUMMARY OF THE INVENTION 
An object, therefore, of the present invention is to provide a new and 
improved projectile release mechanism for facilitating launching a 
self-propelled projectile, particularly a spin-stabilized missile. 
In the exemplary embodiment of the invention, the mechanism includes 
missile support means having rotary means including receptacle means 
defining a spin axis, and fixed support means for supporting the rotary 
means for rotation about the spin axis as well as for movement axially of 
the spin axis. Nozzle means extend from the missile into the receptacle 
means, including fusible joint means for heating by high-temperature 
exhaust gases expelled by the missile to release the missile, and an aft 
nozzle section is movable away from the missile on fusing and separation 
of the fusible joint means. Abutment means are formed on the rotary means 
in the path of movement of the aft nozzle section for striking by the aft 
nozzle section to effect rapid movement of the rotary means and receptacle 
means axially away from the missile following fusing and separation of the 
fusible joint means and on impact of the aft nozzle means with the rotary 
means abutment. 
Preferably, spring biasing means are operatively associated between the 
rotary means and the nozzle means, particularly the aft section of the 
nozzle means, for holding the nozzle means and, thereby, the missile in 
the receptacle means. In addition, detent means are provided between the 
rotary means and the fixed support means to prevent premature motion and 
post separation recoil of the rotary means after striking by the aft 
section of the support means. 
Another feature of the invention includes a register section on the support 
means for receiving the missile, the register section and the missile 
having complementarily engageable, axially spaced concentric land means to 
insure proper alignment of the missile. One of the axially spaced land 
means, the forward land means shown herein, is of a cylindrical 
configuration, and the other or rear land means is of a forwardly opening 
conical configuration. 
Other objects, features and advantages of the invention will be apparent 
from the following detailed description taken in connection with the 
accompanying drawings.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
Referring to the drawings in greater detail, and first to FIG. 1, a 
substantially spherical, spin-stabilized, self-propelled missile 10 is 
shown mounted to the front of a barrel 12 of an assault weapon such as a 
rifle, generally designated 14. The rifle shown is a standard M-16A2 
military rifle or any similar device. 
As shown in FIG. 1 and the enlarged view of FIG. 2, a missile support 
means, generally designated 16, include a front upper attachment portion 
18 with axial motion restraint 19 and a rear attachment portion, generally 
designated 20. Bracket portion 18 is positioned on the barrel 12 whereby 
part of the gas emanating from the barrel is channeled through a 
passageway 22 (FIG. 2) to a firing pin assembly, generally designated 24, 
which is effective to strike a primer on missile 10 to ignite the rocket 
propellant therein, as is known in the art. The latch and the axial motion 
restraint means 19 are provided to lock support means 16 onto the rifle 
barrel. 
Support means 16 also include turbine support portions 28 and 30 which 
support the missile and release mechanism on an axis 32 upwardly inclined 
relative to an intended straight-line path of flight 34 generally parallel 
to the axis of rifle barrel 12. As is known in the art, axis 32 is the 
spin axis of missile 10; i.e. the motor thrust axis of the missile rocket 
motor. Axis 34 which defines the line of flight of the missile is the 
forward velocity component thereof. 
Generally, the self-propelled missile 10 is a spinning projectile launched 
from essentially a zero-length launcher. In other words, this is in 
contrast to a bullet which travels through the entire length of the rifle 
barrel. For accuracy and trajectory repeatability, the missile must be 
maintained in constant alignment with spin axis 32 during spin-up and 
release. Furthermore, since the rifle is fired and recoils during spin-up 
and release of the missile, the missile release must be practically 
instantaneous in order to prevent launcher/projectile impulse moments from 
redirecting the missile during and immediately after release. The present 
invention addresses these problems and has been shown to be effective in 
assuring an undisturbed spin-up and launch event not heretofore available 
with the prior art. 
More particularly, referring specifically to FIG. 2, turbine rotary means, 
generally designated 36, include a plurality of turbine nozzles 38. 
Preferably, four nozzles are provided, 90 degrees apart, to provide 
uniform and equalized torque transmission forces. Four nozzles are used to 
reduce pressure drop variations across the coupling ports and equalize the 
exhaust gas flow through the coupling ports. The multiplicity of turbine 
arms accomplishes this by reducing the back pressure in the turbine plenum 
located between the coupling and the turbine air inlets and allowing 
smoother more evenly distributed exhaust flow through the coupling, 
coupling ports, turbine plenum and turbine arms. In assembly, rotary means 
36 is rotatable within turbine support portions 28 and 30 by appropriate 
bearing means. The rotary means has a forward missile register section 40 
for mating with missile 10, as described hereinafter, an intermediate 
receptacle section 42 journalled in support portion 28, and a rear distal 
end section 44 journalled in support portion 30. Thus, missile register 
section 40, receptacle section 42 and rear distal end section 44 are 
generally coaxial with spin axis 32. 
A nozzle assembly, generally designated 46, includes a fore section 48 and 
an aft section 50 joined by an integral fusible joint means, generally 
designated 52. Intermediate receptacle section 42 of turbine rotary means 
36 forms receptacle means for the nozzle assembly on spin axis 32. The 
fusible joint means 52 is similar to that shown in the aforementioned U.S. 
Pat. No. 4,395,836 and is disposed for heating by high-temperature exhaust 
gases expelled by missile 10 to release the missile from support means 16 
and particularly from rotary turbine means 36. Details of such a fusible 
joint means can be derived from the aforesaid patent which is incorporated 
herein by reference. As disclosed herein, the fusible joint geometry has 
been refined to provide uniform heating and erosion and thereby assure 
that the separation event will occur simultaneously and abruptly across 
the entire joint surface or area, minimizing the time for complete 
separation. More particularly, as seen in FIG. 3, nozzle section 48 is a 
one-piece homogeneous nozzle member. A peripheral ring portion 57 is 
reduced in sectional thickness by appropriate machining operations. A 
precise pattern of equally spaced, axially extending slots or passages 57a 
form the fusible joint which is an integral part of the one-piece nozzle 
member and which is separated by the high-temperature exhaust gases 
expelled by missile 10. The slots, versus the round holes of the 
aforementioned patent, provide an area which is more uniformly heated 
under the influence of a spring preload, and separates more nearly 
instantaneously (e.g. 0.001 seconds) across the whole transverse "plane" 
defined by the slots, with a minimum loss of preload prior to separation. 
On fusing and separation of fusible joint means 52, fore and aft nozzle 
sections 48 and 50, respectively, are completely separated. Once 
separated, aft nozzle section 50 can move rearwardly in the direction of 
arrow "A" within receptacle section 42 of rotary turbine means 36. An 
axial slot 54 in receptacle section 42 guides a pin 56 extending 
therethrough and into aft nozzle section 50, thereby transmitting turbine 
torque to the missile 10. The aft section contains slots that vent exhaust 
gases after separation of the fusible joint. 
The substantially zero-length launching and substantially instantaneous 
release of missile 10 is significantly facilitated by a unique mounting of 
rotary turbine means 36 in support means 16. The missile 10 is seated 
firmly in the register section 40 of rotary means 36 through the agency of 
a coil spring 58, bearing against the termination of rear distal end 
section 44 of the rotary turbine means and a washer 60 fixed to a rod 62 
which, in turn, is fixed to aft nozzle section 50. The forward position of 
rotary means 36 is limited by a snap ring and washer assembly 64 so as to 
properly align the missile's percussion cap with firing pin assembly 24. 
Its aft location is fixed by plunger means 72, described hereinafter. 
Upon fusing or failure of fusible joint means 52, aft nozzle section 50 
recoils rearwardly under the action of the rocket motor gases and the 
preload of spring 58. The aft nozzle section moves rearwardly in the 
direction of arrow "A" and this recoiling mass strikes an abutment seat 66 
on the interior of rotary turbine means 36. The kinetic energy of the 
recoiling mass is transferred to the rotary turbine means and is 
sufficient to overcome the load of spring loaded plunger 72, accelerating 
the rotary turbine means in an aft direction, stripping the mating lands 
(described hereinafter) between register section 40 of the rotary turbine 
means 36 and missile 10, leaving the missile free and with sufficient 
clearance to preclude recontact producing tipoff forces regardless of 
rifle/launcher motions. Rearward movement of the rotary turbine means is 
limited by a shoulder 68 which comes into engagement with the front of 
support means 16, at support portion 28. In essence, the invention 
provides a high speed (microseconds) separation for a more slowly moving 
projectile. In addition, "bounce-back" or recoil of the rotary turbine 
means 36 from impact with the support means is prevented by a detent means 
70 which will project into the path of a ramp latch flange 74. The detent 
means is in the form of a spring loaded plunger 72 which bears against the 
ramped flange 74 about the distal end section 44 of the rotary turbine 
means, thus positioning the turbine in the rearward direction. 
Another feature of the invention includes the provision of complementarily 
engageable, axially spaced concentric land means on register section 40 of 
rotary turbine means 36. More particularly, this interface comprises a 
large-diameter, cylindrical forward land 76 and a small-diameter, shallow 
angle conical aft land 78, both lands being concentric to missile spin 
axis 32. The two axially spaced land means are held in position axially by 
the spring loading (i.e. spring 58) between nozzle 46 and rotary turbine 
means 36, as described above. In contrast to a pair of conical lands, 
cylindrical land means 76 and conical land 78 can be machined within fine 
tolerances to maintain concentricity of the interfacing components and 
also provides a surface for carrying axial loads. Yet, conical land means 
78 still affords the necessary seal for the mechanism. 
The important contribution of the invention in terms of the reduction in 
"tip off" effects and reduction in separation time can best be described 
in relation to FIG. 4. More particularly, referring to that Figure, tip 
off forces "F" cause tip off moments "M". The time interval for which the 
tip off moment is applied to the missile is the moment impulse (Moment 
Impulse=M.multidot.t). The resulting precessional motion is a pointing 
error "E" which is proportional to the moment impulse 
(E=K.multidot.t.multidot.M). In prior constructions, the time "t" in the 
error equation has been relatively large (e.g. 10 milliseconds). This is 
the time during which the projectile has pulled away somewhat from 
register section 40 due to stretch of joint 52 during fusing and when 
"tipoff" forces due to rifle/projectile contact occur, stretch during 
fusing. With the invention, the missile is restrained during coupling 
fusing and, therefore, is not repointed as might be caused by coupling 
fusing initiated tip off forces. With the invention, the aft motion of the 
rear nozzle and its momentum transfer to the rotary means effect rapid aft 
motion of the rotary means. The aft motion effects rapid disengagement 
between the missile and the rotary means (e.g. 50-100 microseconds). This 
reduces both the tipoff impulse and the separation time by a factor of on 
the order of 100 and, thus, reduces the pointing error angle "E" by a 
factor of 100. 
It will be understood that the invention may be embodied in other specific 
forms without departing from the spirit or central characteristics 
thereof. The present examples and embodiments, therefore, are to be 
considered in all respects as illustrative and not restrictive, and the 
invention is not to be limited to the details given herein.