Method and apparatus for aligning spin-stabilized self-propelled missiles

Abstract: For use with a launching apparatus for a spin-stabilized self-propelled missile (10), which includes a rotary missile support (36) defining a spin axis (76) and receiving the missile (10) with a missile axis of rotation (80), and a spring (56) operative to hold the missile (10) on the missile support (36), a method and apparatus are provided for aligning the missile axis (80) of rotation with the spin axis (76). The rotary missile support (36) and missile (10) are supported on a fixture (60) for rotation about the spin axis (76). The rotary missile support (36) and missile (10) then are rotated relative to the fixture (60) about the spin axis (76). The amount of eccentricity (D) between the missile axis (80) of rotation and the spin axis (76) is determined. The missile (10) is restrained for single plane motion and is adjustably moved relative to the rotary support (60) to coincidently align the missile axis (80) of rotation with the spin axis (76). The spring (56) then is tightened to maintain the coincident alignment of the axes.

Description: This invention generally relates to the launching of spin-stabilized self-propelled missiles. 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 term "spherical" herein and in the claims hereof is being used in a generic sense to mean line-of-sight projectiles or missiles. For instance, in the exemplary embodiment herein, the missile is spherical only in the forward half of the missile, the aft half being substantially conical in shape. 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. Such spherical spin-stabilized missiles often 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 achieving missile spin axis alignment 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 propulsion range. Consequently, attempts have been made to provide means for temporarily restraining and automatically releasing a spin-stabilized self-propelled spherical missile during spin-up. Some such attempts are shown in US-A-3245350, US-A-3554078, US-A-4395836 and US-A-4403435, the latter two patents being owned by the present applicant. These patents represent a continuing effort to provide workable spherical spin-stabilized missiles. Generally, a fusible link temporarily restrains and automatically releases the spherical missile during spin-up. Hot missile rocket exhaust gas weakens, by heating, and melts the fusible link which, prior to weakening by softening or melting, secures the missile to a rotary support means. The aforesaid US-A-4395836 shows a novel unitary nozzle member having fusible joint means formed integrally therewith, between the missile and the rotary support means. The aforesaid US-A-4403435 shows an improved nozzle assembly including 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 to improve alignment of the missile with the spin axis during initial separation of the fusible joint means. A somewhat radical departure from the prior art is shown in our copending EP-A-0342294 which is directed to a projectile release mechanism wherein a mass is caused to be urged or propelled rearwardly by the gases of the missile or other suitable stored energy 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. This allows positive missile retention by the launch system rotary means during coupling fusing and therefore eliminates pointing error tip off forces initiated during the coupling fusing process of prior apparatus. That invention represents a vast improvement in the prior art, in that the missile separates from its turbine assembly in less than 0.5 msec., compared to the 10 msec. separation period encountered in earlier designs in the art. This huge reduction in separation time minimizes the transfer of separate impulse forces from the turbine coupling and assembly to the projectile. However, still further problems have been encountered designing such spin-stabilized self-propelled missile systems, a condition which has been termed a "repointing condition" which is caused by projectile construction static and dynamic unbalances. In other words, the axis of the turbine assembly or rotary missile support means attempts to control the system control spin axis, while the relatively large missile segment of the system attempts to create its own axis of rotation. If these axes are not colinear, the missile tends to wobble during spinup and jump to an unaimed heading upon release. For ease of understanding, it is as if a common screwdriver was bent and rapidly rotated when in use. The missile ends up being repointed to an angle generally equal to one-half of the misalignment angle between the axis of the rotary support means (spin axis) and the axis of rotation or geometric axis of the missile. This is what is termed a "repointing condition". Heretofore, attempts to overcome the repointing condition were limited to dynamic balancing of the projectile itself. In other words, weight means were applied to the missile somewhat similar to the common balancing of a vehicle tyre. This approach yielded usable accuracy results, but it is impractical to consider live warhead projectile dynamic balancing in a production environment. The present invention has as its object to provide an alternative to dynamic balancing which can be carried out in a manner compatible with a production environment. This object is met, in accordance with the invention as claimed, by the provision of a method of aligning an axis of rotation of a spin-stabilized self-propelled missile with a spin axis defined by rotary missile support means receiving the missile, the rotary missile support means forming part of a launching device which also includes connection means operative to hold the missile on the support means, the method being characterized by the steps of supporting the rotary missile support means and the missile on a fixture for rotation about the spin axis, rotating the rotary missile support means and the missile relative to the fixture about the spin axis, determining the amount of eccentricity between the missile axis of rotation and the spin axis, adjustably moving the missile relative to the rotary missile support means to coincidentally align the missile axis of rotation with the spin axis, and tightening the connection means to maintain the coincident alignment of the axes. The alignment method of the invention results in a production system whereby the assembled missile and rotary missile support means can be readily mounted on their intended launching device without any other procedures such as static and dynamic balancing being required. The invention also provides an apparatus for performing the alignment method, the apparatus being characterized by a fixture for supporting the rotary missile support means and the missile for rotation about the spin axis, and means on the fixture for determining the amount of eccentricity between the missile axis of rotation and the spin axis whereby the missile can be moved relative to the rotary missile support means while supported on the fixture, to coincidentally align the missile axis of rotation and the spin axis. The fixture may also include side supports for containing the plane of eccentricity and a bottom support for controlling the adjustment motion in the plane of eccentricity. In order that the invention may be well understood there will now be described an embodiment thereof, given by way of example, reference being made to the accompanying drawings, in which like reference numerals identify like elements in the figures, and in which: FIGURE 1 is an elevational view of a spin-stabilized missile mounted on the barrel of a rifle and incorporating a release mechanism or launching apparatus for use with the alignment method and apparatus of the invention; FIGURE 2 is a fragmented side elevational view, partially in section and on an enlarged scale, showing some of the components of the missile and launching apparatus of Figure 1 prior to ignition; FIGURE 3 is a view similar to that of Figure 2, after separation of the fore and aft sections of the nozzle and on impact of the aft section with the launching apparatus; FIGURE 4 is a view similar to that of Figures 2 and 3, showing the turbine assembly driven rearwardly against the launcher; FIGURE 5 is a vertical section through a fixture incorporating the concepts of the invention and for carrying out the method of the invention; and FIGURE 6 is a front elevational view of the fixture of Figure 5. Referring first to Figure 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. The deployment structure may be any fixed or portable structure, and the utility of the invention is not limited to a hand carried weapon such as a rifle. As shown in Figure 1, and the enlarged view of Figure 2, a missile support means, generally designated 16, includes a front upper attachment portion 18 with an axial motion restraint means, generally designated 19. The attachment portion 18 is generally tubular for positioning over the barrel 12, and a tightening screw 20 fixes the attachment portion to the barrel. A nut 21 locks the axial restraint means 19 in place by retaining a clamp bar 19a. The attachment portion 18 is positioned on the barrel 12, whereby part of the gas emanating from the barrel is channelled through a passageway 22 (Fig. 2) to a firing pin assembly, generally designated 24, which is effective to strike a primer on the missile 10 to ignite the rocket propellant therein, as is known in the art. The support means 16 also include turbine support land portions 28 and 30 (Fig. 3) 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 35 of the rifle barrel 12. As is known in the art, the axis 32 is the spin axis of the missile and turbine assembly (described hereinafter); i.e., the motor thrust axis of the missile rocket motor. The axis 34 which defines the line of flight of the missile is the forward velocity or down range 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 or launch tube. For accuracy and trajectory repeatability, the missile must be maintained in constant alignment with the 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 the release process. These problems are addressed in the aforesaid copending application Serial No. 195,657 which is incorporated herein by reference. That invention has been shown to be effective in assuring an undisturbed spin-up and launch event superior to any prior art and, as stated above, the missile disengages in less than 0.5 msec. Suffice it to say herein, and still referring specifically to Figure 2, a rotary missile support means or turbine rotary assembly, generally designated 36, includes a plurality of turbine nozzles 38. Preferably, four nozzles are provided, 90# apart, to provide uniform and equalized torque transmission forces. The rotary missile support means 36 has annular registration surfaces 39a and 39b for registering with complementary registration surfaces on the missile 10. In assembly, the rotary missile support means 36 includes land portions 40 and 42 for precisely registering with the complementary land portions 28 and 30, respectively, on the support means 16. These land portions are concentric with the spin axis 32. A nozzle assembly, generally designated 46, includes a fore section 48 and an aft section 50 fixed to a rearwardly projecting bolt-like shaft 52 having an externally threaded rear end. A meltable joint 53 integrally joins the fore and aft sections 48 and 50, respectively. The rotary missile support means 36 has an internal, radially inwardly projecting annular flange 54. A support or connection means in the form of a coil spring 56 is sandwiched between the flange 54 and a tightening nut 58 threaded onto the rear end of the shaft 52. Therefore, the missile 10 and the nozzle assembly 46 are held within the missile rotary support means 36 by the spring 56 and the nut 58. In other words, the rotary missile support means 36 provides receptacle means for the missile 10 and the nozzle assembly 46 to support the missile and nozzle assembly on the spin axis 32. Very briefly, referring to Figure 3, when the meltable joint 53 separates, the aft section 50 of the nozzle assembly 46 is driven aftwardly in the direction of an arrow "X" until it strikes the turbine assembly 36 at shoulders 59. The turbine assembly 36 then is driven aftwardly in the direction of arrows "Y" as shown in Figure 4 until it is stopped by shoulders 61 on a locking collar "C". As amplified to considerable extent heretofore, further problems exist when the axis of rotation of the missile 10 is not colinear with the spin axis 32. This can result from manufacturing tolerances, expansion and contraction allowances for temperature variances and other variables during manufacture whereby clearances result in the interfaces, such as the registration surfaces 39a and 39b between the missile 10 and the rotary support means 36, as well as the interfaces between the nozzle assembly 46 and the rotary support means 36. As a result, a "repointing condition" may occur during separation should the manufacturing assembly process cause the axis of rotation of the missile 10 to vary from the spin axis 32 of the rotary support means 36, as allowed by the component design and fabrication tolerances. In order to alleviate these problems, and referring to Figure 3, the invention contemplates a producibility system of aligning the missile and nozzle assembly with respect to the rotary support means prior to assembly on the intended launching device such as support means 16 on the rifle 14. To this end, the invention contemplates the provision of a fixture, generally designated 60 (Fig. 5), which has annular lands 62 and 64 precisely simulating the lands 28 and 30 of the actual support means 16 of the intended launching device. The fixture 60 has a base portion 68 (or any support structure) for rigidly supporting the fixture on an appropriate support structure 70. Frame portions 72 and 74 project upwardly from the base 68 and terminate in and define the annular lands 62 and 64, respectively. This rigid construction defines a spin axis 76 simulating the spin axis 32 of the rotary support means 36 when mounted on the rifle support means 16 as described in relation to Figures 1 and 2. A frame arm 77 projects upwardly from the land 62 and then forwardly and outwardly over an area where a missile 10 would be disposed when mounted in the fixture, as described below. A measuring gauge, generally designated 78, is mounted on the forward distal end of the frame arm 77 for quantifying missile and rotary means axes eccentricities. As indicated in Figure 5, the missile 10 has a geometric axis 80 which runs from a geometric front centre point 81 rearwardly through the centre of gravity of the missile and through the axial centre of the nozzle 46. Ideally, the missile axis 80 defines the axis of rotation of the missile 10 when in flight and, ideally, should coincide with or be colinear with the spin axis 76. However, most likely, the missile axis of rotation 80 and the spin axis 76 will be out of alignment when the missile 10 and the nozzle 46 are positioned within the rotary missile support means 36, as indicated by the slight angle represented by arrows "D" (Fig. 3). As stated above, this misalignment results in a "repointing condition" during missile separation as the missile 10 attempts to jump off the spin axis 76 and results in a "wobbling" effect during flight. Heretofore, the axes were aligned by dynamic balancing, similar to balancing an ordinary automobile tyre and wheel, which does not lend itself to a practical production environment. This invention is directed to solving these problems and providing a capability of producibly controlling the alignment of the missile axis of rotation 80 and the spin axis 76, as in the fixture 60, prior to mounting the missile 10 in its appropriate launching device such as the support means 16 on the rifle 14. Specifically, the method generally comprises the steps of providing a fixture, such as the fixture 60, for receiving the rotary missile support means 36, with the missile 10 and the nozzle means 46 positioned within the receptacle means defined by the support means, and supporting the rotary missile support means for rotation about its spin axis, i.e., the axis 76. In other words, the lands 40 and 42 of the rotary support means 36 are positioned in the lands 62 and 64, respectively, of the fixture 60. However, it should be noted that the fixture 60 requires only two point contact at the land area (64) and two point contact at the land area (62) as long as the aft points are above the spin axis 76 and the forward points are below the spin axis. Gravity loading with the missile 10 in place will then provide the spin axis 76 spacial defunction. Still generally, the rotary missile support means 36, with the missile 10 and the nozzle means 46 positioned therein, then are rotated about the axis 76 while positioned in the fixture 60. At this point, the spring 56 is in the fully collapsed condition. As the rotary missile support means 36, the nozzle means 46 and the missile 10 are rotated, the amount of eccentricity between the missile axis of rotation 80 and the spin axis 76 is determined. This determination is made by the gauge 78 which has a projecting head 82 located at the maximum diameter of the missile 10 taken generally perpendicular to its axis of rotation 80. The amount of eccentricity, or what is commonly termed "run out", of the missile axis and the spin axis can be determined by the amount the missile will move the head 82 of the gauge 78 away from the axis 76. For instance, as the missile 10 rotates, its "run out" away from the axis 76 will move the head 82 upwardly in the direction of arrow "E". More particularly, referring to Figure 6 in conjunction with Figure 5, the fixture 60 has a pair of L-shaped side supports 84 slidably mounted on the base 68 by appropriate bracket means 86 fixed to the base, and including any appropriate means, such as set screws 88 threaded in the bracket means, for securing the side supports in any position of adjustment relative to the base 68. These side supports (84) are provided for containing the plane of eccentricity of the missile 10. Referring back to Figure 5, an adjustable bottom support, generally designated 90, is provided for controlling the adjusting motion of the missile 10 in the plane of eccentricity confined by the side supports 84. The bottom support 90 includes a fixed block 92 secured to the base 68 and including a rearwardly inclined ramp surface 94 on top of the fixed block. An adjusting block 96 has a screw 98 extending therethrough and threaded through an upwardly projecting flange portion 100 of the fixed block 92. The adjusting block 96 has a bottom angled cam surface 102 which rides up the ramp surface 94 of the fixed block 92. With this construction, rotation of the screw 98 will cause the adjusting block 96 to move in the direction of a double-headed arrow "X" whereby the adjusting block 96 can move vertically for engaging the underside of the missile 10 and thereby providing a bottom support for the missile. With the side supports 84 in an "open" position as shown in phantom in Figure 6, and with the adjusting block 96 moved forwardly (to the left in Fig. 5), the invention contemplates a procedure for aligning the missile 10 in the rotary support means 36 as now described. First, the missile 10 is loaded into the rotary support means 36, and the missile/support means assembly is loaded into the fixture 60 as described above, with the spring 56 in collapsed condition. The missile/support means assembly then is rotated to determine a plane of its high point and its low point in relation to a vertical direction. This best can be understood with reference to Figure 6 wherein the "high point" of the missile 10 is shown in full lines and the "low point" of the missile is shown in phantom, thereby defining a vertical plane of eccentricity designated 104. In other words, the assembly should be rotated so that its high point (of the missile) is at the top/dead-centre of the plane 104. These high and low positions easily can be determined by the gauge 78, with the head 82 riding on the top of the missile. The head 82 may be on a distal end of a plunger 106 (Fig. 5) which is effective to rotate an indicating needle 108 relative to a dial 110. Once the high point and low point of the missile 10 are determined by using the gauge 78, the missile is rotated so that it is at its "low point" as indicated by the dial 110 and generally represented in phantom in Figure 6. The side supports 84 then are closed inwardly until they contact the sides of the missile 10 to contain and restrain the missile horizontally relative to the fixture 60. The side supports 84 are locked in place relative to the base 68 by the set screws 88. The adjusting block 98 (Fig. 5) of the bottom support 90 then is moved up the ramp surface 94 until it contacts the bottom of the missile 10. The spring 56 then is loosened and the adjusting block 96 is moved further to move the missile 10 upwardly to a point half-way between the previously determined high and low points. The spring 56 then is tightened. The side and bottom supports 84, 90 then are moved away from the missile 10 and the missile is rotated to test for any eccentricity. If any eccentricity still appears by reading movement of the needle 108 of the gauge 78, the above steps can be repeated as required.

Claims: A method of aligning an axis of rotation of a spin-stabilized self-propelled missile (10) with a spin axis (76) defined by rotary missile support means (36) receiving the missile (10), the rotary missile support means (36) forming part of a launching device (16) which also includes connection means (56, 58) operative to hold the missile (10) on the support means (36), the method being characterized by the steps of supporting the rotary missile support means (36) and the missile (10) on a fixture (60) for rotation about the spin axis (76), rotating the rotary missile support means (36) and the missile (10) relative to the fixture (60) about the spin axis (76), determining the amount of eccentricity between the missile axis of rotation (80) and the spin axis (76), adjustably moving the missile (10) relative to the rotary missile support means (36) to coincidentally align the missile axis of rotation (80) with the spin axis (76), and tightening the connection means (56, 58) to maintain the coincident alignment of the axes (76, 80). A method as claimed in claim 1, wherein the rotary missile support means (36) has land means (40, 42) concentric with its spin axis (76) for mating with complementary land means (28, 30) on its intended launching device (16), and including providing the fixture (60) with land means (62, 64) simulating the complementary land means (28, 30) on the launching device (16), and the supporting step comprises supporting the rotary missile support means (36) on the fixture (60) with their respective land means (40, 42; 28, 30) mated. A method as claimed in claim 1 or claim 2, wherein the determining step is carried out along a line generally perpendicular to the spin axis (76). A method as claimed in any of the preceding claims, wherein the missile (10) is substantially spherical, and the determining step is carried out at a diameter of the missile (10) perpendicular to the missile axis of rotation (80). A method as claimed in any of the preceding claims, including repeating the determining step and the moving step before tightening the connection means (56, 58). A method as claimed in any of the preceding claims, including restraining the missile (10) on the fixture (60) in a plane (104) of maximum eccentricity before moving the missile (10) relative to the rotary missile support means (36). A method as claimed in any of the preceding claims, wherein the rotary missile support means (36) has receptacle means (36) defining the spin axis (76), the missile (10) has nozzle means (38) extending into the receptacle means (36) and defining the missile axis of rotation (80), and the connection means (56, 58) is operatively associated between the rotary missile support means (36) and the nozzle means (38) for holding the nozzle means (38) and, thereby, the missile (10) in the receptacle means (36), wherein the method further comprises the step of the fixture (60) receiving the rotary missile support means (36) with the missile (10) and the nozzle means (38) positioned in the receptacle means (36), the fixture (60)supporting the support means (36) for rotation about its spin axis (76). An apparatus for aligning an axis of rotation (80) of a spin-stabilized self-propelled missile (10) with a spin axis (76) defined by rotary missile support means (36) receiving the missile (10), the rotary missile support means (36) forming part of a launching device (16) which also includes connecting means (56, 58) operative to hold the missile (10) on the support means (36), the apparatus being characterized by a fixture (60) for supporting the rotary missile support means (36) and the missile (10) for rotation about the spin axis (76), and means (78) on the fixture (60) for determining the amount of eccentricity between the missile axis of rotation (80) and the spin axis (76) whereby the missile (10) can be moved relative to the rotary missile support means (36), while supported on the fixture (60), to coincidentally align the missile axis of rotation (80) and the spin axis (76). An apparatus as claimed in claim 8, wherein the rotary missile support means (36) for the missile (10) has land means (40, 42) concentric with its spin axis (76) for mating with complementary land means (28, 30) on its intended launching device (16), and the fixture (60) includes land means (62, 64) simulating the complementary land means (28, 30) on the launching device (16) for supporting the rotary missile support means (36) on the fixture (60) with their respective land means (40, 42; 28, 30) mated. An apparatus as claimed in claim 8 or claim 9, wherein the means (78) for determining is located on a diameter of the missile (10) perpendicular to the missile axis of rotation (80). An apparatus as claimed in any of claims 8 to 10, wherein the means (78) for determining is located on a line generally perpendicular to the spin axis (76). An apparatus as claimed in any of claims 8 to 11, including means (84) for constraining the missile (10) in a plane (104) of its maximum eccentricity.

Inventors: STEELE MICHAEL F, STEELE, MICHAEL F.
Applicants: BRUNSWICK CORP, BRUNSWICK CORPORATION
Application year: 1991
Main classification: F42B35
IPC codes: F42B35, F41F3