Title of Invention	"A DEPLOYMENT MECHANISM FOR A MISSILE HAVING AERODYNAMIC FIN"
Abstract	At least one aerodynamic fin (312) and a deployment mechanism (314) for a missile (10) having the at least one aerodynamic fin (312), the deployment mechanism (314) comprising a tubular cam (334) having at least one cam slot (354); a cam pin (336) connected to the fin (312) and extending into the at least one cam slot (354); and a spring (38/138/238/338) mountable in a missile (10) for deploying the fin from a stowed orientation to a deployed orientation that is different from the stowed orientation; and characterized in that the fin (312) has a connecting portion (350) with a spherical attachment point (351) manufactured to fit with a very close tolerance against the inner diameter of the tubular cam (334).

Title of Invention

"A DEPLOYMENT MECHANISM FOR A MISSILE HAVING AERODYNAMIC FIN"

Abstract

At least one aerodynamic fin (312) and a deployment mechanism (314) for a missile (10) having the at least one aerodynamic fin (312), the deployment mechanism (314) comprising a tubular cam (334) having at least one cam slot (354); a cam pin (336) connected to the fin (312) and extending into the at least one cam slot (354); and a spring (38/138/238/338) mountable in a missile (10) for deploying the fin from a stowed orientation to a deployed orientation that is different from the stowed orientation; and characterized in that the fin (312) has a connecting portion (350) with a spherical attachment point (351) manufactured to fit with a very close tolerance against the inner diameter of the tubular cam (334).

Full Text	The present invention relates to deployment mechanism for aerodynamic fins. The present invention generally relates to ordnance having stowable fins, and, more particularly, to a missile having a deployment mechanism for deploying the fins. Many types of ordnance utilize two or more protruding surfaces to affect the fluid flow around the ordnance, thereby facilitating control of its trajectory toward a target. Exemplary types of such ordnance include missiles, rockets, bombs, torpedoes and the like. For example, missiles generally have an approximately cylindrical body, with at least two aerodynamic surfaces or fins that extend outwardly from the sides of the missile body to affect the aerodynamic characteristics of the missile in flight. The fins typically have an airfoil shape that is oriented edge-on or slightly inclined relative to the airflow when the missile is flying in a straight line. These fins may be, for example, static (fixed) or dynamic (selectively movable, i.e., controllable). Fixed fins generally are used to stabilize the missile during flight and do not move once fully deployed. Controllable fins (control fins) are used to control or steer the missile by selectively varying the attitude of the fins relative to the airflow under the direction of the missile's control system. In many cases, the fins are stowed in a position adjacent the outside surface of or within the missile body during storage and mounting on a vehicle prior to use. In some cases, the missile is stored in a tube, canister or other protective casing, and the protective casing also may serve as a launch tube. The fins are stowed to reduce the effective diameter of the missile, permitting more missiles to be stored and/or transported in a limited space. It also reduces the likelihood of damage to the fins during storage and handling. Additionally, it allows for the maximum use of the internal space of the missile for electronic components and warheads. The fins are extended from the stowed position shortly after deployment of the missile, either during mounting or launch of the missile. Various relatively complex deployment mechanisms have been developed to permit the fins to be stowed, deployed and locked into place. Control fins may further be moved (usually only rotated) by an actuator system once the control fins are deployed. The mechanisms presently used to deploy the fins tend to be relatively heavy, complex and expensive to design, build and maintain. Moreover, some mechanisms occupy a relatively large volume within the missile, a significant disadvantage because of the limited space within the missile. SUMMARY OF THE INVENTION There is a need for a simple and reliable device to support, deploy and lock stowable ordnance fins into a deployed configuration. The present invention provides a deployment mechanism for deploying stowable fins that meets this need and provides further advantages in cost, weight and space savings. More particularly, the present invention provides a missile with the deployment mechanism that automatically deploys a fin from a stowed orientation to a deployed orientation as soon as the fin is released. The deployment mechanism includes a spring that provides a biasing force that urges the fin to move quickly, simply and reliably from the stowed orientation to the deployed orientation. The deployment mechanism also includes one or more cam slots or other means for guiding the fin from the stowed orientation to the deployed orientation. An exemplary deployment mechanism for the missile includes a tubular cam body that can be mounted in a cylindrical cavity in the missile body. A drive pin is connected to the cam body through the spring which biases the drive pin to the deployed orientation. The fin is connected to a cam pin that extends into cam slots in the cam body to guide the fin as it is deployed. The cam pin also interconnects the fin and the drive pin. The drive pin and the spring thus cooperate to move the fin from the stowed orientation to the deployed orientation, while the cam pin and the cam slots guide the fin as it is deployed. Thecam slots may also rotate the fin as it is deployed and/or lock the fin in place. Such a deployment mechanism can be used with either a fixed fin or a dynamic control fin, in any type of ordnance having stowable fins, including the missile described herein. To simplify the description, reference herein is specifically directed to missiles, but such reference includes other types of ordnance where the description would be applicable. More particularly, one aspect of the invention relates to a deployment mechanism for a missile having at least one aerodynamic fin. The deployment mechanism comprises a spring mountable in a missile for deploying the at least one fin. The deployment mechanism is operable to move the at least one fin from a stowed orientation to a deployed orientation that is different from the stowed orientation. Another aspect of the invention relates to the deployment mechanism further including a tubular cam having at least one cam slot and a cam pin connected to the at least one fin. The spring is connected to the cam pin to urge the cam pin to a deployed configuration. The deployed configuration includes the at least one fin in the deployed orientation. The cam pin is movable along and guided by the at least one cam slot to pivot the at least one fin and to rotate the at least one fin from the stowed orientation to the deployed orientation. To the accomplishment of the foregoing and related ends, the invention provides the features hereinafter fully described and particularly pointed out in the claims. The following description and the annexed drawings set forth in detail certain illustrative embodiments of the invention. These embodiments are indicative, however, of but a few of the various ways in which the principles of the invention may be employed. Other objects, advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the drawings. DRAWINGS Figure 1 is a partial and schematic perspective view of a forward section of an exemplary missile body with aerodynamic fins in a stowed configuration. Figure 2 is a partial and schematic perspective view of the missile shown in Figure 1 with the fins in a deployed configuration. Figure 3 is a schematic cross-sectional view of a section of the missile body showing the fin and a sectioned deployment mechanism in accordance with the invention in the stowed configuration. Figure 4 is a schematic cross-sectional view of a section of the missile body showing the fin and the sectioned deployment mechanism in the deployed configuration. Figure 5 is an elevational view of a tubular cam in accordance with the invention. Figure 6 is an exploded schematic perspective view of the fin and the deployment mechanism in accordance with another embodiment of the invention. Figure 7 is a partial and schematic perspective view of the fin and the deployment mechanism of the embodiment shown in Figure 6 in the stowed configuration partially in section. Figure 8 is a partial and schematic perspective view of the fin and the deployment mechanism of the embodiment shown in Figure 6 in the deployed configuration partially in section. Figure 9 is a partial and schematic cross-sectional view of a fin locking mechanism provided by the present invention. Figures 10a- 10e are a sequence of schematic perspective views of the fin and the deployment mechanism shown in Figure 6 transitioning from the stowed configuration to the deployed configuration in accordance with the invention. Figures 1 1a-1 lb are schematic perspective views of a tubular cam in accordance with yet another embodiment of the invention. Figure 12 is a partial and schematic cross-sectional view of the fin and the deployment mechanism shown in Figures 10a-10b in an actuator system of the missile. Figure 13 is an exploded schematic perspective view of the fin and the deployment mechanism in accordance with still another embodiment of the invention. Figure 14 is an exploded schematic perspective view of the fin and the deployment mechanism shown in Figure 13 from a different angle. Figure 15 is a schematic bottom view of the fin shown in Figure 13. In the detailed description that follows, similar components in different embodiments will have a similar reference numeral incremented by 100. For example, in a first embodiment, a cam is assigned reference number 34. Subsequent embodiments may use reference numbers 134, 234, 334, etc., for the cam bodies of subsequent embodiment, although the cam body may have a different configuration in the different embodiments. For the sake of brevity, in-depth descriptions of similar components may be omitted from descriptions of subsequent embodiments. DETAILED DESCRIPTION Referring now to the drawings, and initially to Figures 1 and 2, the present invention provides ordnance, such as a missile 10, having a plurality of fins 12 for stabilizing or controlling the missile during flight. The fins 12 include at least one stowable fin 12 and a deployment mechanism 14 for moving the fin 12 from a stowed configuration (Figure 1) to a deployed configuration (Figure 2) so that the missile 10 can be stored or launched in a more compact configuration. The illustrated missile 10 has four fins 12 mounted to a generally cylindrical body (missile body) 16 having a longitudinal axis 18. Although the present description refers to the missile 10 shown in the drawings, the illustrated missile 10 represents any type of ordnance that uses stowable fins and is not limited to a missile. Each fin 12 has a leading edge 20 and a trailing edge 22 that bound the width of the fin 12, and a longitudinal axis 24 that extends approximately along the length of the fin 12. The leading edge 20 of the fin 12 preferably faces in a forward direction generally toward the leading or forward end of the missile 10 during flight. The thickness of the fin 12 is less than its width or length, and the geometry of the fin 12 is selected for its intended application. In the stowed configuration shown in Figure 1, the fins 12 lie adjacent to a surface 26 of the missile body 16. The longitudinal axis 24 of each fin 12 approximately parallels the longitudinal axis 18 of the missile body 16, and the leading edge 20 and the trailing edge 22 of each fin 12 face sideways to provide a compact stowed configuration wherein the missile 10 occupies a minimum volume. In the illustrated embodiment, the missile body 16 has a longitudinally extending recess 28 (Figure 2) in its surface 26 for receiving the fin 12 in the stowed or stored configuration. With the fin 12 stowed and received in the recess 28, an outer surface 30 (Figure 1) of the fin 12 generally conforms to the outer surface 26 of the missile 10. The recess 28 has a shape and size sufficient to receive the fin 12 while minimizing the volume of the missile 10 taken up by the recess 28. In the illustrated embodiment, the recess 28 extends from an end of the fin 12 that is attached to the missile 10 toward the forward end of the missile 10. In the deployed configuration shown in Figure 2, each fin 12 extends from the surface of the missile body 16. The longitudinal axis 24 of the fin 12 is approximately perpendicular to the longitudinal axis 18 of the missile body 16, and the leading edge 20 generally faces toward the forward end of the missile 10. The fin 12 is connected to the missile body 16 through the deployment mechanism 14, which moves the fin 12 from the stowed orientation to the deployed orientation. Referring now to Figures 3-5, an assembly, including the fin 12 and the deployment mechanism 14, is mounted at least partially in a cavity 32 in the missile body 16 (Figures 3-4). The deployment mechanism 14 includes a tubular cam 34, a cam pin 36, a drive spring 38, and a drive pin 40. The cam 34 has an internal step, shelf or ledge 42 formed by an abrupt change in its internal diameter for engaging an outer coil 44 of the drive spring 38, which in the illustrated embodiment is a conical spring. An inner coil 46 of the drive spring 38 is connected to the drive pin 40 for applying force thereto. In the illustrated embodiment, the inner coil 46 of the drive spring 38 engages a flange portion 48 of the drive pin 40 that has a greater lateral extent than an adjacent portion of the drive pin 40. In other words, the flange portion 48 is an annular ring or disk at one end of a smaller diameter (generally cylindrical) portion of the drive pin 40. The drive spring 38 is mounted inside the cam 34, interposed between the shelf 42 and the flange portion 48 of the drive pin 40 to urge or bias the drive pin 40 to the deployed orientation. The drive pin 40 interconnects the drive spring 38 and the cam pin 36. In the illustrated embodiment, a connecting portion 50 of the fin 12 has a central notch 52 at a free end thereof and the cam pin 36 is mounted to traverse the central notch 52. The end portions of the cam pin 36 extend beyond the edges of the connecting portion 50 to engage cam slots 54. The drive pin 40 is connected to the cam pin 36 within the central notch 52. The cam pin 36 is rotatable with respect to at least one of the drive pin 40 and the connecting portion 50 of the fin 12 to allow the fin 12 to pivot about a longitudinal axis of the cam pin 36. The cam pin 36 also rotates about a central axis approximately coextensive with a longitudinal axis 56 of the cam 34. The cam pin 36 generally remains perpendicular to the longitudinal axis 56 of the cam 34 as it rotates. The cam pin 36 is guided by at least one cam slot or groove 54 extending from an inner surface 58 of the cam 34 that receives and guides end portions of the cam pin 36. In other words, the cam pin 36 acts as a follower as it traces the cam slots 54. The cam slots 54 may extend partially or completely through the wall of the cam 34. In the illustrated embodiment, the cam 34 has a pair of diametrically opposed and approximately helical slots 54 that guide the cam pin 36 to simultaneously rotate and translate along the longitudinal axis 56 of the cam 34 (Figure 5). The shape of the cam slots 54 may be tailored to vary the path and orientation of the fin 12 as the cam pin 36 moves between the stored and deployed configurations. The cam 34 guides the deployment of the fin 12 and generally is fixed in the cavity 32 against rotation in at least one direction, for example, by mating a threaded end (mounting end 60, Figure 5) of the cam 34 with corresponding threads in the cavity 32 (not shown). This helps to keep the cam 34 from coming loose as the fin 12 rotates into position. An opposite end of the cylindrical cam 34 (a working end 62), includes a pair of stepped faces 64 and 66 (hereinafter pivot face 64 and stop face 66) separated by two laterally spaced upright faces (one shown, Figure 5) 68, extending generally parallel to the longitudinal axis 56 of the cam 34. The upright faces 68 are interposed between the pivot face 64 at the lower step and the stop face 66 at an upper step. The pivot face 64 is formed by the absence of a semi-cylindrical section at the working end 62 of the cam 34. The cam 34 is mounted to the missile 10 such that the pivot face 64 is even with or proud of the surface of the recess 28 adjacent the cavity 32. The stop face 66 generally extends above the missile surface 26. As the fin 12 is moved from the stowed orientation to the deployed orientation, the fin 12 simultaneously pivots about the pivot face 64 and rotates about the longitudinal axis 56 of the cam 34, with an end 72 of the fin 12 engaging the stop face 66 in the deployed orientation. The laterally extending end portions of the cam pin 36 travel through the cam slots 54 until the cam pin 36 reaches the deployed configuration (Figure 2) with the lateral end portions at or near the respective ends of the cam slots 54. The end portions of the cam slots 54 may provide positive stops for the cam pin 36 corresponding to the stored and deployed orientations of the fin 12. In other words, the cam pin 36 may engage the ends of the cam slots 54 at the stored and deployed orientations of the fin 12, respectively. In operation, the cam slots 54 effect simultaneous rotational and pivotal movement of the fin 12 in response to the telescoping axial movement of the drive pin 40. Retraction of the drive pin 40 by the drive spring 38 urges the cam pin 36 (in the illustrated orientation) through the cam slots 54 simultaneously rotating the cam pin 36 and the fin 12 through approximately ninety degrees (90°) from the stowed orientation (Figure 3) to the deployed orientation (Figure 4). At the same time, the connecting portion 50 of the fin 12 pivots about the pivot face 64 of the cam 34 and moves into the cam 34. The pivot face 64 effectively functions as a fulcrum for moving the longitudinal axis 24 of the fin 12 as the fin 12 moves from an orientation substantially parallel to the longitudinal axis 18 of the missile body 16 (Figure 3) to an orientation substantially perpendicular to the longitudinal axis 18 of the missile body 16 (Figure 4). Stated another way, the cam pin 36 and the cam slots 54 translate the axial movement of the drive pin 40 into both a rotational and axial movement of the fin 12 as the cam pin 36 follows the cam slots 54. With the fin in the stowed orientation (Figure 3), the drive spring 38 stores potential energy. When released, the deployment mechanism 14 simultaneously pivots and rotates the fin 12 from the stowed orientation (Figure 3) to the deployed orientation (Figure 4). The energy of the drive spring 38 drives the cam pin 36 along the longitudinal axis 56 of the cam 34 and also holds the fin 12 in the deployed orientation once deployed. Resistance created by airflow over the missile 10 also may help to deploy and to retain the fin 12 in the deployed orientation. The assembly can, of course, be modified to accommodate different sizes, configurations and types of ordnance. For example, the drive springs 38 are selected to provide the appropriate power for the size of the fins 12. A locking mechanism (not shown) may further be provided to retain the fin 12 in the deployed orientation. For example, the end portions of the cam pin 36 may be spring-loaded and outwardly biased into blind rather than through slots, and a locking detent (not shown) may be provided at an end of the cam slots 54. The spring-loaded portions would travel along the cam slots 54 until reaching respective detents, where the end portions would extend further into the detents to lock the cam pin 36 in place. Alternatively, a bump (not shown) may be formed in the cam slots 54 over which the spring-loaded end portions would readily pass over in a first direction, but which would inhibit or prevent the spring-loaded end portions from passing in a second direction opposite the first direction. A retaining mechanism (not shown) also may be used to prevent the fins 12 from moving prematurely from the stowed orientation. For example, a tab on the fin 12 may be held in place by a flange extending from the outer surface 26 of the missile body 16 to help hold the fin 12 in the stowed orientation until deployed. Locking pins (not shown) also may be used. Turning to Figures 6-10, another assembly of a fin 112 and an alternative deployment mechanism 114 is shown. To facilitate the description, similar elements have been given similar reference numbers incremented by a factor of one hundred (100). As in the prior embodiment, the deployment mechanism 114 includes a cam 134, a cam pin 136, a drive spring 138 and a pivot pin 140. The cam pin 136 spans a central notch 152 in a connecting portion 150 of the fin 112 and extends into a cam slot 154 in the wall of the cam 134. In this embodiment, the relative positions of the drive spring 38 (Figure 3) and the drive pin 40 (Figure 3) of the prior embodiment have been reversed. Consequently, the drive spring 138 is interposed between the cam pin 136 and the pivot pin 140 and does not directly act on the cam body 134. The drive spring 138 is an extension spring having a loop or hook 174 at one end for engaging the cam pin 136 and a bent tab 176 at the opposite end. The pivot pin 140 in turn is held in a disk 178 at the mounting end of the cam 134. The disk 178 may be secured to the cam 134 by corresponding threads (not shown) on the disk 178 and at the mounting end of the cam 134. Alternatively, the disk 178 may be held against an internal shelf 142 of the cam 134 (Figure 8) by the drive spring 138. The cam 134 includes the internal shelf 142 that forms a stop that limits how far the disk 178 can extend into the cam 134. The drive spring 138 holds the pivot pin 140 in the disk 178. However, the pivot pin 140 is rotatable relative to the disk 178 about a longitudinal axis generally parallel to a longitudinal axis 156 of the cam 134 as the drive spring 138 rotates with the cam pin 136. This arrangement further reduces the number of moving parts. Further, this arrangement provides additional force on the cam pin 136 which increases the reliability of the deployment mechanism 114. Further still, this arrangement reduces the number of assembly steps, for example, by allowing the tab 176 of the drive spring 138 to be inserted into the pivot pin 140 from the outside of the cam 134. Turning to a detailed description of individual components, the disk 178 has a large diameter ring portion 180 and a small diameter disk portion 182 adjacent the ring portion 180. The disk portion 182 fits inside the cam 134 and engages the internal shelf 142 when the disk 178 is fully tightened or inserted. The disk portion 182 also includes a hole or slot or other opening 184 for receiving the pivot pin 140 extending therethrough as will be explained below. The disk portion 182 is connected to an inner diameter of the ring portion 180 thereby forming a cavity inside the ring portion 180 for receiving the pivot pin 140. The pivot pin 140 is similar to the drive pin 40 shown in Figure 3. The pivot pin 140 has a generally cylindrical body 186 having a through hole 188 extending transverse to the longitudinal axis of the body for receipt of the tab portion 176 of the drive spring 138. A flange portion 148 having a greater lateral extent is connected to an adjacent portion of the cylindrical body 186. In the illustrated embodiment, the flange portion 148 is an annular ring or disk having a diameter that is larger than the opening 184 in the disk portion 182 of the disk 178. When the pivot pin 140 is inserted through the opening in the disk 178, the flange portion 148 is received in the cavity. When assembled, the pivot pin 140 is free to rotate about a longitudinal axis corresponding to the longitudinal axis 156 of the cam 134. During the deployment motion, the pivot pin 140 rotates with the drive spring 138 as the drive spring 138 rotates with the cam pin 136. The drive spring 13 8 generally extends along a longitudinal axis perpendicular to the cam pin 136 and is telescopically received in the tubular cam 134 for extension and retraction generally parallel to the longitudinal axis 156 of the cam 134. The drive spring 138 is an extension spring formed of several coils. On one end, the last coil forms the hook 174. On the other end, the last coil is formed into the tab 176. The pivot pin 140 and the disk 178 anchor the drive spring 138 to the cam 134. The drive spring 138 interconnects the pivot pin 140 and the cam pin 136 to pull the cam pin 136 through the cam slots 154 and toward the pivot pin 140. The cam pin 136 interconnects the drive spring 138 and the fin 112. In the illustrated embodiment, the cam pin 136 has an annular groove 190 for receiving the hook portion 174 of the drive spring 138 within the central notch 152 of the fin 112. The annular groove 190 inhibits lateral motion of the hook 174 relative to the cam pin 136. In the illustrated embodiment, respective ends of the cam slot 154 extend in a direction substantially parallel to the longitudinal axis 156 of the cam 134 to prevent rotation of the fin 112 when the cam pin 136 is moving through that portion of the cam slot 154. Accordingly, the cam slot 154 forces the fin 112 to pivot from the stowed orientation without rotating right away, unlike the previous embodiment. At an upper or working end 162 of the cam 134, the cam 134 has a central notch or axially relieved portion 164 formed between two laterally spaced wall sections 168 and 170. A wedge block 192 (Figure 6) is formed on the axially relieved portion 164 of the cam 134 between the wall sections 168 and 170. The wedge 192 is located approximately in the center of the axially relieved portion 164 and provides a fulcrum or pivot point upon which the fin 112 initially pivots as it deploys. The wedge 192 also may be used as a stop to further prevent or minimize the fin 112 from rocking when it is in the deployed orientation. A rocking motion of the fin 112 may occur in a direction toward and away from the forward end of the missile. The wedge 192 has a narrow stop on top that engages the fin 112 during deployment. The wedge 192 has a wide base to distribute the stresses acting upon it. From the axially relieved portion 164, the wall section 170 includes a ramp 194 that spirals downward, toward the opposite end of the cam 134, in a clockwise direction. The ramp 194 has a slope that helps to control the fin 112 as it is deployed. As the fin 112 is deployed, the end or base 172 of the fin 112 engages the ramp 194 and spirals down the slope until the fin 112 engages a stop 196 (Figure 9) formed by an end of the opposing wall section 168. The wall section 168 generally has a uniform height that extends above the lower end of the ramp 194 and prevents further rotation of the fin 112. When the fin 112 engages the stop 196, the stop 196 prevents further rotation of the fin 112, but allows the fin 112 to move parallel to the longitudinal axis 156 of the cam 134 as will be further explained below. In the illustrated embodiment, the fin 112 has a tapered tab 198 formed therein at the base of the fin 112 to help lock the fin 112 in the deployed orientation. The cam 134 further includes a slot 200 between the end of the ramp 194 and the stop 196. The slot 200 forms part of a fin locking mechanism 202. Referring additionally to Figure 9, the tapered tab 198 may have a raised rim 204 on a lower end thereof, the tapered tab 198 engages the fin locking mechanism 202 when the fin 112 is in the deployed configuration. The tapered tab 198 is shaped to slide into the slot 200 in a first direction, downward in the illustrated orientation, but would be inhibited or prevented from passing in a second direction opposite the first direction by the raised rim 204. The raised rim 204 engages a corresponding raised stop 206 portion of the fin locking mechanism 202 and thus prevents the fin 112 from moving from the deployed orientation. To assemble the deployment mechanism 114, the drive spring 138 is inserted into the cam 134. The tab 176 of the drive spring 138 is inserted through the hole 184 and into the through hole 188 of the pivot pin 140. The pivot pin 140 is inserted into the disk 178. The connecting portion 150 of the fin 112 is inserted into the cam 134, the hook 174 of the drive spring 138 is placed within the notch 152 and the cam pin 136 is inserted through the connecting portion 150 and within the hook 174 of the drive spring 138 through the slots 154. Thus, the hook 174 of the drive spring 138 is placed in the annular groove 190 of the cam pin 136 and within the notch 152 of the connecting portion 150 of the fin 112. The disk 178 is secured in the cam 134 by the spring 138. Sequential images illustrating the deployment of the fin 112 from the stowed orientation to the deployed orientation are shown in Figures 10a-lOe. The fin 112 is shown in the stowed orientation in Figure 10a. As soon as the fin 112 is released, the fin 112 pivots about the wedge 192 of the axially relieved portion 164 of the cam 134. The fin 112 then pivots approximately ninety degrees (901) as the cam pin 136 moves within the cam slots 154 in an axial direction towards the disk 178. Next, the laterally extending end portions of the cam pin 136 spiral through the cam slots 154 (M2). The fin 112 simultaneously rotates with the cam pin 136 and moves downward into the cam 134 with the cam pin 136 (M2). The end 172 of the fin 112 engages and slides along the ramp 194 of the wall section 170 until the end 172 engages the stop 196 of the wall section 168 (M2). Next, the fin 112 moves in an axial direction towards the disk 178 (M3). The tapered tab 198 of the fin 112 engages the fin locking mechanism 202 as the end portions of the cam pin 136 follow the end portions 208 of the slots 154. The forward end of the fin 112 engages the stop of the wedge 192. Thus, the fin 112 is fully deployed with a leading edge 120 facing the forward end of the missile 10 (Figure 2). The fin locking mechanism 202 cooperates with the end portions 208 of the cam slots 154 and the stop of the wedge 192 to reduce the rocking of the fin 112 relative to the cam 134 during the remainder of the missile's flight. Specifically, the wedge 192 prevents the fin 112 from coming out of the locking mechanism 202 during a forward rocking motion of the fin 112. The deployment mechanism 114 shown in Figures 6-10 is continuously active as is the case with the deployment mechanism 14 shown in Figures 3 and 4. In other words, the deployment mechanism 114 continuously applies a force to the fins 112. This urges the fins 112 to rotate from the stowed orientation to the deployed orientation. During the assembly of the missile, the fins 112 are assembled in or moved to the stowed orientation and placed inside a missile launch tube, for example (not shown). As a result of placing the fins 112 in the stowed orientation, the deployment mechanism 114 continuously applies a force to the pivot pin 140 along the longitudinal axis 156 of the cam 134 toward the disk 178. Without a locking mechanism to retain the fins 112 against the missile body 16 (Figure 1), the fins 112 pivot about the axially relieved portion 164 with the distal end of the fins 112 moving away from the surface of the missile 26 (Figure 1) and engaging an inner surface of the launch tube. The inner surface of the launch tube thus prevents the fins 112 from fully deploying. During launch, the distal ends of the fins 112 engage the inner surface of the launch tube as the missile moves down the launch tube. Once the fins 112 clear the end of the launch tube, the deployment mechanisms 114 can complete the deployment of the fins 112. The drive springs 138 urge the laterally extending end portions of cam pins 136 to move through the cam slots 154. The fins 112 pivot and then rotate with the cam pins 136 until the bases of the fins 112 engage the fin locking mechanisms 202 and the stops of the wedges 192 of the cams 134. Thus, the fins 112 fully deploy with the leading edges 120 facing the forward end of the missile 10 (Figure 1) and with a longitudinal axis 124 of each fin 112 extending substantially perpendicular to the surface of the missile 26 (Figure 2). In an alternative embodiment, the deployment mechanism 114 may be manually or automatically activated. A retaining mechanism (not shown), such as a retaining pin, may be used to hold each fin 112 in the stowed orientation. Once the retaining pin is removed, the deployment mechanism 114 deploys the fin 112 as described in the preceding paragraph. Figures lla-llband 12 show another assembly ofa fin 212 and another embodiment of a deployment mechanism 214. The deployment mechanism 214 is substantially the same as the previously described deployment mechanism 114 (Figure 6). However, the deployment mechanism 214 includes an alternative cam 234. In this embodiment, the disk 178 (Figure 6) in the previous embodiment is incorporated into the mounting end of the cam 234 to form a single unit. In other words, the cam 234 has a closed end 278 that performs the function of the disk 178 (Figure 6). The closed end 278 is in the shape ofa disk and has an opening 284 therethrough. The opening 284 may be shaped as two interconnecting openings with a large diameter opening 285 near an outer edge of the closed end 278 and a small diameter opening 287 near the center of the closed end 278. Surrounding the small diameter opening 287 is a recessed surface 289 for receiving the flange 248 of the pivot pin 240. The closed end 278 of the cam 234 allows the final assembly to be completed completely from the exterior. This embodiment further reduces the number of parts of the deployment mechanism 214. The assembly, including the control fin 212 and the deployment mechanism 214 is shown in combination with an actuator 291 in a deployed configuration in Figure 12. In this embodiment, the cam 234 functions as an actuator shaft rotatably mounted to the actuator 291 for selectively rotating the control fin 212 about a longitudinal axis 256 of the cam 234 once the control fin 212 is in the deployed orientation. A missile guidance controller (not shown) selectively controls the actuator 291 to rotate the control fin 212 relative to the direction of airflow for controlled flight of the missile. More specifically, as shown in Figure 12, the cam 234 is seated in the actuator 291 within an upper bearing 293 and a lower bearing 295. The cam 234 has threads on an outer surface of the lower end for receiving a threaded nut 297 thereon. The cam 234 also has an upper land or ridge 299. The upper ridge 299 engages the inner race of the upper bearing 293, and the nut 297 engages an inner race of the lower bearing 295. As the nut 297 is tightened and torqued, the two bearings 293 and 295 are trapped across a mounting block 301 of the actuator 291 and pre-loaded to secure the cam 234 to the actuator 291. This keeps the cam 234 from rattling around and allows the actuator 291 to rotate the cam 234, and thus the fin 212, at high speeds. Now referring to Figures 13-15, yet another assembly is shown. This assembly includes a fin 312 and a deployment mechanism 314. The fin 312 has a connecting portion 350 with a spherical attachment point 351. The spherical attachment point 351 has a central notch 352, which separates the spherical attachment point 351 into two generally hemispherical portions. The spherical attachment point 351 also has a through hole 353 for receiving a cam pin 336 therein. The spherical attachment point 351 is manufactured to fit with a very close tolerance against the inner diameter of the cam 334. This allows the spherical attachment point 351 to reduce the stress on the cam pin 336 as the fin 312 pivots and rotates from the stowed orientation to the deployed orientation. In particular, the spherical attachment point 351 reduces the stresses acting on the cam pin 336 in the fully deployed orientation of the fin 212 by transferring those stresses to the spherical attachment point 351. At a base 372 of the fin 312, wedge shape protrusions extend from opposite faces of the fin 312 to form a key 398. The key 398 cooperates with the deployment mechanism 314 to help hold the fin 312 in the deployed orientation as will be clear from the following explanation. The deployment mechanism 314 is substantially similar to the previously described deployment mechanism 114 (Figure 6) except as particularly described in the following paragraphs. The deployment mechanism 314 includes the cam 334, the cam pin 336, a drive spring 338, a pivot pin 340 and a disk 378 assembled as described with respect to Figures 6-10. The cam 334 has a relieved portion 364 and two laterally spaced upright sections 368 and 370. Between the laterally spaced upright sections 368 and 370 and opposite the relieved portion 364 is a keyway 355. The keyway 355 provides additional stability for the fin 312 upon full deployment and prevents or minimizes rocking of the fin 312 during the remainder of the missile's flight. The invention thus provides a simple and reliable mechanism to both hold the fins in a stowed position and to release the fins to a deployed configuration. Further, no parts of the device are shed or broken away upon deployment of the fins, thereby minimizing or eliminating the risk of injury to the launch vehicle or operator. Although the invention has been shown and described with respect to certain preferred embodiments, it is obvious that equivalent alterations and modifications will occur to others skilled in the art upon the reading and understanding of this specification and the annexed drawings. In particular regard to the various functions performed by the above described components (assemblies, devices, sensors, circuits, etc.), the terms (including a reference to a Ameans@) used to describe such components are intended to correspond, unless otherwise indicated, to any component which performs the specified function of the described component (i.e., that is functionally equivalent), even though not structurally equivalent to the disclosed structure which performs the function in the herein illustrated exemplary embodiments of the invention. In addition, while a particular feature of the invention may have been disclosed with respect to only one of several embodiments, such feature may be combined with one or more other features of the other embodiments as may be desired and advantageous for any given or particular application. We claim 1. At least one aerodynamic fin (312) and a deployment mechanism (314) for a missile (10) having the at least one aerodynamic fin (312), the deployment mechanism (314) comprising: a tubular cam (334) having at least one cam slot (354); a cam pin (336) connected to the fin (312) and extending into the at least one cam slot (354); and a spring (38/138/238/338) mountable in a missile (10) for deploying the fin from a stowed orientation to a deployed orientation that is different from the stowed orientation; and characterized in that the fin (312) has a connecting portion (350) with a spherical attachment point (351) manufactured to fit with a very close tolerance against the inner diameter of the tubular cam (334). 2. The fin (312) and the deployment mechanism (314) as claimed in claim 1, wherein the spherical attachment point (351) has a central notch (352) which separates the spherical attachment point (351) into two generally hemispherical portions. 3. The fin (312) and the deployment mechanism (314) as claimed in claim 2, wherein the spherical attachment point (351) also has a through hole (353) for receiving the cam pin (336) therein. 4. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 3, wherein the spring (338) is an extension spring. 5. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 4, wherein the spring (338) is connected to the cam pin (336) to urge the cam pin (336) to a deployed configuration in which the fin (312) is in the deployed orientation, and wherein the cam pin (336) is movable along and guided by the at least one cam slot (354) to pivot the fin (312) and to rotate the fin (312) from the stowed orientation to the deployed orientation. 6. The fin (312) and deployment mechanism (311) as claimed in any of claims 1 to 5, wherein the tubular cam (334) has an upper face (62) that forms a fulcrum about which the fin (312) pivots. 7. The aerodynamic fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 6, wherein the fin (312) includes a protruding key (398) adjacent an end thereof and the deployment mechanism (314) includes a keyway (355) for receiving the key to inhibit a rocking movement of the fin. 8. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 7, wherein at the base of the fin (312), wedge shape protrusions extend from opposite faces of the fin (312) to form the key (398). 9. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 8, wherein the cam (334) has a relieved portion (364) and two laterally spaced upright sections (368, 370) and wherein the keyway (355) is between the laterally spaced upright sections (368, 370) and opposite the relieved portion (364). 10. A missile (10) comprising the aerodynamic fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 9. 11. A missile (10) as claimed in any of claims 1 to 10, comprising cylindrical missile body (16) having an outer surface (26) and a recess (28) in the surface sized to receive the fin (312) in the stowed orientation. 12. A missile (10) as claimed in claim 10 or claim 11, comprising an actuator (291) for selectively rotating the control fin (312) about a longitudinal axis (356) once the control fin (312) is in the deployed orientation. 13. A missile (10) as claimed in claim 12, wherein the actuator (291) rotates the fin (312) relative to the direction of airflow for controlled flight of the missile.

Full Text

The present invention relates to deployment mechanism for aerodynamic fins.
The present invention generally relates to ordnance having stowable fins, and, more particularly, to a missile having a deployment mechanism for deploying the fins.
Many types of ordnance utilize two or more protruding surfaces to affect the fluid flow around the ordnance, thereby facilitating control of its trajectory toward a target. Exemplary types of such ordnance include missiles, rockets, bombs, torpedoes and the like. For example, missiles generally have an approximately cylindrical body, with at least two aerodynamic surfaces or fins that extend outwardly from the sides of the missile body to affect the aerodynamic characteristics of the missile in flight. The fins typically have an airfoil shape that is oriented edge-on or slightly inclined relative to the airflow when the missile is flying in a straight line. These fins may be, for example, static (fixed) or dynamic (selectively movable, i.e., controllable). Fixed fins generally are used to stabilize the missile during flight and do not move once fully deployed. Controllable fins (control fins) are used to control or steer the missile by selectively varying the attitude of the fins relative to the airflow under the direction of the missile's control system.
In many cases, the fins are stowed in a position adjacent the outside surface of or within the missile body during storage and mounting on a vehicle prior to use. In some cases, the missile is stored in a tube, canister or other protective casing, and the protective casing also may serve as a launch tube. The fins are stowed to reduce the effective diameter of the missile, permitting more missiles to be stored and/or transported in a limited space. It also reduces the likelihood of damage to the fins during storage and handling. Additionally, it allows for the maximum use of the internal space of the missile for electronic components and warheads.
The fins are extended from the stowed position shortly after deployment of the missile, either during mounting or launch of the missile. Various relatively complex deployment mechanisms have been developed to permit the fins to be stowed, deployed and locked into place. Control fins may further be moved (usually only rotated) by an actuator system once the control fins are deployed.
The mechanisms presently used to deploy the fins tend to be relatively heavy, complex and expensive to design, build and maintain. Moreover, some mechanisms occupy
a relatively large volume within the missile, a significant disadvantage because of the limited space within the missile.
SUMMARY OF THE INVENTION
There is a need for a simple and reliable device to support, deploy and lock stowable ordnance fins into a deployed configuration. The present invention provides a deployment mechanism for deploying stowable fins that meets this need and provides further advantages in cost, weight and space savings.
More particularly, the present invention provides a missile with the deployment mechanism that automatically deploys a fin from a stowed orientation to a deployed orientation as soon as the fin is released. The deployment mechanism includes a spring that provides a biasing force that urges the fin to move quickly, simply and reliably from the stowed orientation to the deployed orientation. The deployment mechanism also includes one or more cam slots or other means for guiding the fin from the stowed orientation to the deployed orientation.
An exemplary deployment mechanism for the missile includes a tubular cam body that can be mounted in a cylindrical cavity in the missile body. A drive pin is connected to the cam body through the spring which biases the drive pin to the deployed orientation. The fin is connected to a cam pin that extends into cam slots in the cam body to guide the fin as it is deployed. The cam pin also interconnects the fin and the drive pin. The drive pin and the spring thus cooperate to move the fin from the stowed orientation to the deployed orientation, while the cam pin and the cam slots guide the fin as it is deployed. Thecam slots may also rotate the fin as it is deployed and/or lock the fin in place. Such a deployment mechanism can be used with either a fixed fin or a dynamic control fin, in any type of ordnance having stowable fins, including the missile described herein. To simplify the description, reference herein is specifically directed to missiles, but such reference includes other types of ordnance where the description would be applicable.
More particularly, one aspect of the invention relates to a deployment mechanism for a missile having at least one aerodynamic fin. The deployment mechanism comprises a spring mountable in a missile for deploying the at least one fin. The deployment mechanism is operable to move the at least one fin from a stowed orientation to a deployed orientation that is different from the stowed orientation.
Another aspect of the invention relates to the deployment mechanism further including a tubular cam having at least one cam slot and a cam pin connected to the at least one fin. The spring is connected to the cam pin to urge the cam pin to a deployed configuration. The deployed configuration includes the at least one fin in the deployed orientation. The cam pin is movable along and guided by the at least one cam slot to pivot the at least one fin and to rotate the at least one fin from the stowed orientation to the deployed orientation.
To the accomplishment of the foregoing and related ends, the invention provides the features hereinafter fully described and particularly pointed out in the claims. The following description and the annexed drawings set forth in detail certain illustrative embodiments of the invention. These embodiments are indicative, however, of but a few of the various ways in which the principles of the invention may be employed. Other objects, advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the drawings.
DRAWINGS
Figure 1 is a partial and schematic perspective view of a forward section of an exemplary missile body with aerodynamic fins in a stowed configuration.
Figure 2 is a partial and schematic perspective view of the missile shown in Figure 1 with the fins in a deployed configuration.
Figure 3 is a schematic cross-sectional view of a section of the missile body showing the fin and a sectioned deployment mechanism in accordance with the invention in the stowed configuration.
Figure 4 is a schematic cross-sectional view of a section of the missile body showing the fin and the sectioned deployment mechanism in the deployed configuration.
Figure 5 is an elevational view of a tubular cam in accordance with the invention.
Figure 6 is an exploded schematic perspective view of the fin and the deployment mechanism in accordance with another embodiment of the invention.
Figure 7 is a partial and schematic perspective view of the fin and the deployment mechanism of the embodiment shown in Figure 6 in the stowed configuration partially in section.
Figure 8 is a partial and schematic perspective view of the fin and the deployment mechanism of the embodiment shown in Figure 6 in the deployed configuration partially in section.
Figure 9 is a partial and schematic cross-sectional view of a fin locking mechanism provided by the present invention.
Figures 10a- 10e are a sequence of schematic perspective views of the fin and the deployment mechanism shown in Figure 6 transitioning from the stowed configuration to the deployed configuration in accordance with the invention.
Figures 1 1a-1 lb are schematic perspective views of a tubular cam in accordance with yet another embodiment of the invention.
Figure 12 is a partial and schematic cross-sectional view of the fin and the deployment mechanism shown in Figures 10a-10b in an actuator system of the missile.
Figure 13 is an exploded schematic perspective view of the fin and the deployment mechanism in accordance with still another embodiment of the invention.
Figure 14 is an exploded schematic perspective view of the fin and the deployment mechanism shown in Figure 13 from a different angle.
Figure 15 is a schematic bottom view of the fin shown in Figure 13.
In the detailed description that follows, similar components in different embodiments will have a similar reference numeral incremented by 100. For example, in a first embodiment, a cam is assigned reference number 34. Subsequent embodiments may use reference numbers 134, 234, 334, etc., for the cam bodies of subsequent embodiment, although the cam body may have a different configuration in the different embodiments. For the sake of brevity, in-depth descriptions of similar components may be omitted from descriptions of subsequent embodiments.
DETAILED DESCRIPTION
Referring now to the drawings, and initially to Figures 1 and 2, the present invention provides ordnance, such as a missile 10, having a plurality of fins 12 for stabilizing or controlling the missile during flight. The fins 12 include at least one stowable fin 12 and a deployment mechanism 14 for moving the fin 12 from a stowed configuration (Figure 1) to a deployed configuration (Figure 2) so that the missile 10 can be stored or launched in a more compact configuration. The illustrated missile 10 has four fins 12 mounted to a generally
cylindrical body (missile body) 16 having a longitudinal axis 18. Although the present description refers to the missile 10 shown in the drawings, the illustrated missile 10 represents any type of ordnance that uses stowable fins and is not limited to a missile.
Each fin 12 has a leading edge 20 and a trailing edge 22 that bound the width of the fin 12, and a longitudinal axis 24 that extends approximately along the length of the fin 12. The leading edge 20 of the fin 12 preferably faces in a forward direction generally toward the leading or forward end of the missile 10 during flight. The thickness of the fin 12 is less than its width or length, and the geometry of the fin 12 is selected for its intended application.
In the stowed configuration shown in Figure 1, the fins 12 lie adjacent to a surface 26 of the missile body 16. The longitudinal axis 24 of each fin 12 approximately parallels the longitudinal axis 18 of the missile body 16, and the leading edge 20 and the trailing edge 22 of each fin 12 face sideways to provide a compact stowed configuration wherein the missile 10 occupies a minimum volume. In the illustrated embodiment, the missile body 16 has a longitudinally extending recess 28 (Figure 2) in its surface 26 for receiving the fin 12 in the stowed or stored configuration. With the fin 12 stowed and received in the recess 28, an outer surface 30 (Figure 1) of the fin 12 generally conforms to the outer surface 26 of the missile 10. The recess 28 has a shape and size sufficient to receive the fin 12 while minimizing the volume of the missile 10 taken up by the recess 28. In the illustrated embodiment, the recess 28 extends from an end of the fin 12 that is attached to the missile 10 toward the forward end of the missile 10.
In the deployed configuration shown in Figure 2, each fin 12 extends from the surface of the missile body 16. The longitudinal axis 24 of the fin 12 is approximately perpendicular to the longitudinal axis 18 of the missile body 16, and the leading edge 20 generally faces toward the forward end of the missile 10. The fin 12 is connected to the missile body 16 through the deployment mechanism 14, which moves the fin 12 from the stowed orientation to the deployed orientation.
Referring now to Figures 3-5, an assembly, including the fin 12 and the deployment mechanism 14, is mounted at least partially in a cavity 32 in the missile body 16 (Figures 3-4). The deployment mechanism 14 includes a tubular cam 34, a cam pin 36, a drive spring 38, and a drive pin 40. The cam 34 has an internal step, shelf or ledge 42 formed by an
abrupt change in its internal diameter for engaging an outer coil 44 of the drive spring 38, which in the illustrated embodiment is a conical spring. An inner coil 46 of the drive spring 38 is connected to the drive pin 40 for applying force thereto. In the illustrated embodiment, the inner coil 46 of the drive spring 38 engages a flange portion 48 of the drive pin 40 that has a greater lateral extent than an adjacent portion of the drive pin 40. In other words, the flange portion 48 is an annular ring or disk at one end of a smaller diameter (generally cylindrical) portion of the drive pin 40. The drive spring 38 is mounted inside the cam 34, interposed between the shelf 42 and the flange portion 48 of the drive pin 40 to urge or bias the drive pin 40 to the deployed orientation.
The drive pin 40 interconnects the drive spring 38 and the cam pin 36. In the illustrated embodiment, a connecting portion 50 of the fin 12 has a central notch 52 at a free end thereof and the cam pin 36 is mounted to traverse the central notch 52. The end portions of the cam pin 36 extend beyond the edges of the connecting portion 50 to engage cam slots 54. The drive pin 40 is connected to the cam pin 36 within the central notch 52. The cam pin 36 is rotatable with respect to at least one of the drive pin 40 and the connecting portion 50 of the fin 12 to allow the fin 12 to pivot about a longitudinal axis of the cam pin 36. The cam pin 36 also rotates about a central axis approximately coextensive with a longitudinal axis 56 of the cam 34. The cam pin 36 generally remains perpendicular to the longitudinal axis 56 of the cam 34 as it rotates.
The cam pin 36 is guided by at least one cam slot or groove 54 extending from an inner surface 58 of the cam 34 that receives and guides end portions of the cam pin 36. In other words, the cam pin 36 acts as a follower as it traces the cam slots 54.
The cam slots 54 may extend partially or completely through the wall of the cam 34. In the illustrated embodiment, the cam 34 has a pair of diametrically opposed and approximately helical slots 54 that guide the cam pin 36 to simultaneously rotate and translate along the longitudinal axis 56 of the cam 34 (Figure 5). The shape of the cam slots 54 may be tailored to vary the path and orientation of the fin 12 as the cam pin 36 moves between the stored and deployed configurations.
The cam 34 guides the deployment of the fin 12 and generally is fixed in the cavity
32 against rotation in at least one direction, for example, by mating a threaded end
(mounting end 60, Figure 5) of the cam 34 with corresponding threads in the cavity 32 (not
shown). This helps to keep the cam 34 from coming loose as the fin 12 rotates into position. An opposite end of the cylindrical cam 34 (a working end 62), includes a pair of stepped faces 64 and 66 (hereinafter pivot face 64 and stop face 66) separated by two laterally spaced upright faces (one shown, Figure 5) 68, extending generally parallel to the longitudinal axis 56 of the cam 34. The upright faces 68 are interposed between the pivot face 64 at the lower step and the stop face 66 at an upper step. The pivot face 64 is formed by the absence of a semi-cylindrical section at the working end 62 of the cam 34. The cam 34 is mounted to the missile 10 such that the pivot face 64 is even with or proud of the surface of the recess 28 adjacent the cavity 32. The stop face 66 generally extends above the missile surface 26.
As the fin 12 is moved from the stowed orientation to the deployed orientation, the fin 12 simultaneously pivots about the pivot face 64 and rotates about the longitudinal axis 56 of the cam 34, with an end 72 of the fin 12 engaging the stop face 66 in the deployed orientation. The laterally extending end portions of the cam pin 36 travel through the cam slots 54 until the cam pin 36 reaches the deployed configuration (Figure 2) with the lateral end portions at or near the respective ends of the cam slots 54. The end portions of the cam slots 54 may provide positive stops for the cam pin 36 corresponding to the stored and deployed orientations of the fin 12. In other words, the cam pin 36 may engage the ends of the cam slots 54 at the stored and deployed orientations of the fin 12, respectively.
In operation, the cam slots 54 effect simultaneous rotational and pivotal movement of
the fin 12 in response to the telescoping axial movement of the drive pin 40. Retraction of
the drive pin 40 by the drive spring 38 urges the cam pin 36 (in the illustrated orientation)
through the cam slots 54 simultaneously rotating the cam pin 36 and the fin 12 through
approximately ninety degrees (90°) from the stowed orientation (Figure 3) to the deployed
orientation (Figure 4). At the same time, the connecting portion 50 of the fin 12 pivots about
the pivot face 64 of the cam 34 and moves into the cam 34. The pivot face 64 effectively
functions as a fulcrum for moving the longitudinal axis 24 of the fin 12 as the fin 12 moves
from an orientation substantially parallel to the longitudinal axis 18 of the missile body 16
(Figure 3) to an orientation substantially perpendicular to the longitudinal axis 18 of the
missile body 16 (Figure 4). Stated another way, the cam pin 36 and the cam slots 54
translate the axial movement of the drive pin 40 into both a rotational and axial movement of
the fin 12 as the cam pin 36 follows the cam slots 54.
With the fin in the stowed orientation (Figure 3), the drive spring 38 stores potential energy. When released, the deployment mechanism 14 simultaneously pivots and rotates the fin 12 from the stowed orientation (Figure 3) to the deployed orientation (Figure 4). The energy of the drive spring 38 drives the cam pin 36 along the longitudinal axis 56 of the cam 34 and also holds the fin 12 in the deployed orientation once deployed. Resistance created by airflow over the missile 10 also may help to deploy and to retain the fin 12 in the deployed orientation. The assembly can, of course, be modified to accommodate different sizes, configurations and types of ordnance. For example, the drive springs 38 are selected to provide the appropriate power for the size of the fins 12.
A locking mechanism (not shown) may further be provided to retain the fin 12 in the deployed orientation. For example, the end portions of the cam pin 36 may be spring-loaded and outwardly biased into blind rather than through slots, and a locking detent (not shown) may be provided at an end of the cam slots 54. The spring-loaded portions would travel along the cam slots 54 until reaching respective detents, where the end portions would extend further into the detents to lock the cam pin 36 in place. Alternatively, a bump (not shown) may be formed in the cam slots 54 over which the spring-loaded end portions would readily pass over in a first direction, but which would inhibit or prevent the spring-loaded end portions from passing in a second direction opposite the first direction.
A retaining mechanism (not shown) also may be used to prevent the fins 12 from moving prematurely from the stowed orientation. For example, a tab on the fin 12 may be held in place by a flange extending from the outer surface 26 of the missile body 16 to help hold the fin 12 in the stowed orientation until deployed. Locking pins (not shown) also may be used.
Turning to Figures 6-10, another assembly of a fin 112 and an alternative deployment mechanism 114 is shown. To facilitate the description, similar elements have been given similar reference numbers incremented by a factor of one hundred (100). As in the prior embodiment, the deployment mechanism 114 includes a cam 134, a cam pin 136, a drive spring 138 and a pivot pin 140. The cam pin 136 spans a central notch 152 in a connecting portion 150 of the fin 112 and extends into a cam slot 154 in the wall of the cam 134. In this embodiment, the relative positions of the drive spring 38 (Figure 3) and the drive pin 40 (Figure 3) of the prior embodiment have been reversed. Consequently, the drive spring 138
is interposed between the cam pin 136 and the pivot pin 140 and does not directly act on the cam body 134.
The drive spring 138 is an extension spring having a loop or hook 174 at one end for engaging the cam pin 136 and a bent tab 176 at the opposite end. The pivot pin 140 in turn is held in a disk 178 at the mounting end of the cam 134. The disk 178 may be secured to the cam 134 by corresponding threads (not shown) on the disk 178 and at the mounting end of the cam 134.
Alternatively, the disk 178 may be held against an internal shelf 142 of the cam 134 (Figure 8) by the drive spring 138. The cam 134 includes the internal shelf 142 that forms a stop that limits how far the disk 178 can extend into the cam 134. The drive spring 138 holds the pivot pin 140 in the disk 178. However, the pivot pin 140 is rotatable relative to the disk 178 about a longitudinal axis generally parallel to a longitudinal axis 156 of the cam 134 as the drive spring 138 rotates with the cam pin 136. This arrangement further reduces the number of moving parts. Further, this arrangement provides additional force on the cam pin 136 which increases the reliability of the deployment mechanism 114. Further still, this arrangement reduces the number of assembly steps, for example, by allowing the tab 176 of the drive spring 138 to be inserted into the pivot pin 140 from the outside of the cam 134.
Turning to a detailed description of individual components, the disk 178 has a large diameter ring portion 180 and a small diameter disk portion 182 adjacent the ring portion 180. The disk portion 182 fits inside the cam 134 and engages the internal shelf 142 when the disk 178 is fully tightened or inserted. The disk portion 182 also includes a hole or slot or other opening 184 for receiving the pivot pin 140 extending therethrough as will be explained below. The disk portion 182 is connected to an inner diameter of the ring portion 180 thereby forming a cavity inside the ring portion 180 for receiving the pivot pin 140.
The pivot pin 140 is similar to the drive pin 40 shown in Figure 3. The pivot pin 140
has a generally cylindrical body 186 having a through hole 188 extending transverse to the
longitudinal axis of the body for receipt of the tab portion 176 of the drive spring 138. A
flange portion 148 having a greater lateral extent is connected to an adjacent portion of the
cylindrical body 186. In the illustrated embodiment, the flange portion 148 is an annular
ring or disk having a diameter that is larger than the opening 184 in the disk portion 182 of
the disk 178. When the pivot pin 140 is inserted through the opening in the disk 178, the
flange portion 148 is received in the cavity. When assembled, the pivot pin 140 is free to rotate about a longitudinal axis corresponding to the longitudinal axis 156 of the cam 134. During the deployment motion, the pivot pin 140 rotates with the drive spring 138 as the drive spring 138 rotates with the cam pin 136.
The drive spring 13 8 generally extends along a longitudinal axis perpendicular to the cam pin 136 and is telescopically received in the tubular cam 134 for extension and retraction generally parallel to the longitudinal axis 156 of the cam 134. The drive spring 138 is an extension spring formed of several coils. On one end, the last coil forms the hook 174. On the other end, the last coil is formed into the tab 176.
The pivot pin 140 and the disk 178 anchor the drive spring 138 to the cam 134. The drive spring 138 interconnects the pivot pin 140 and the cam pin 136 to pull the cam pin 136 through the cam slots 154 and toward the pivot pin 140. The cam pin 136 interconnects the drive spring 138 and the fin 112. In the illustrated embodiment, the cam pin 136 has an annular groove 190 for receiving the hook portion 174 of the drive spring 138 within the central notch 152 of the fin 112. The annular groove 190 inhibits lateral motion of the hook 174 relative to the cam pin 136.
In the illustrated embodiment, respective ends of the cam slot 154 extend in a direction substantially parallel to the longitudinal axis 156 of the cam 134 to prevent rotation of the fin 112 when the cam pin 136 is moving through that portion of the cam slot 154. Accordingly, the cam slot 154 forces the fin 112 to pivot from the stowed orientation without rotating right away, unlike the previous embodiment.
At an upper or working end 162 of the cam 134, the cam 134 has a central notch or
axially relieved portion 164 formed between two laterally spaced wall sections 168 and 170.
A wedge block 192 (Figure 6) is formed on the axially relieved portion 164 of the cam 134
between the wall sections 168 and 170. The wedge 192 is located approximately in the
center of the axially relieved portion 164 and provides a fulcrum or pivot point upon which
the fin 112 initially pivots as it deploys. The wedge 192 also may be used as a stop to
further prevent or minimize the fin 112 from rocking when it is in the deployed orientation.
A rocking motion of the fin 112 may occur in a direction toward and away from the forward
end of the missile. The wedge 192 has a narrow stop on top that engages the fin 112 during
deployment. The wedge 192 has a wide base to distribute the stresses acting upon it.
From the axially relieved portion 164, the wall section 170 includes a ramp 194 that spirals downward, toward the opposite end of the cam 134, in a clockwise direction. The ramp 194 has a slope that helps to control the fin 112 as it is deployed. As the fin 112 is deployed, the end or base 172 of the fin 112 engages the ramp 194 and spirals down the slope until the fin 112 engages a stop 196 (Figure 9) formed by an end of the opposing wall section 168. The wall section 168 generally has a uniform height that extends above the lower end of the ramp 194 and prevents further rotation of the fin 112. When the fin 112 engages the stop 196, the stop 196 prevents further rotation of the fin 112, but allows the fin 112 to move parallel to the longitudinal axis 156 of the cam 134 as will be further explained below.
In the illustrated embodiment, the fin 112 has a tapered tab 198 formed therein at the base of the fin 112 to help lock the fin 112 in the deployed orientation. The cam 134 further includes a slot 200 between the end of the ramp 194 and the stop 196. The slot 200 forms part of a fin locking mechanism 202.
Referring additionally to Figure 9, the tapered tab 198 may have a raised rim 204 on a lower end thereof, the tapered tab 198 engages the fin locking mechanism 202 when the fin 112 is in the deployed configuration. The tapered tab 198 is shaped to slide into the slot 200 in a first direction, downward in the illustrated orientation, but would be inhibited or prevented from passing in a second direction opposite the first direction by the raised rim 204. The raised rim 204 engages a corresponding raised stop 206 portion of the fin locking mechanism 202 and thus prevents the fin 112 from moving from the deployed orientation.
To assemble the deployment mechanism 114, the drive spring 138 is inserted into the cam 134. The tab 176 of the drive spring 138 is inserted through the hole 184 and into the through hole 188 of the pivot pin 140. The pivot pin 140 is inserted into the disk 178. The connecting portion 150 of the fin 112 is inserted into the cam 134, the hook 174 of the drive spring 138 is placed within the notch 152 and the cam pin 136 is inserted through the connecting portion 150 and within the hook 174 of the drive spring 138 through the slots 154. Thus, the hook 174 of the drive spring 138 is placed in the annular groove 190 of the cam pin 136 and within the notch 152 of the connecting portion 150 of the fin 112. The disk 178 is secured in the cam 134 by the spring 138.
Sequential images illustrating the deployment of the fin 112 from the stowed orientation to the deployed orientation are shown in Figures 10a-lOe. The fin 112 is shown in the stowed orientation in Figure 10a. As soon as the fin 112 is released, the fin 112 pivots about the wedge 192 of the axially relieved portion 164 of the cam 134. The fin 112 then pivots approximately ninety degrees (901) as the cam pin 136 moves within the cam slots 154 in an axial direction towards the disk 178.
Next, the laterally extending end portions of the cam pin 136 spiral through the cam
slots 154 (M2). The fin 112 simultaneously rotates with the cam pin 136 and moves
downward into the cam 134 with the cam pin 136 (M2). The end 172 of the fin 112 engages
and slides along the ramp 194 of the wall section 170 until the end 172 engages the stop 196
of the wall section 168 (M2). Next, the fin 112 moves in an axial direction towards
the disk 178 (M3). The tapered tab 198 of the fin 112 engages the fin locking mechanism 202 as the end portions of the cam pin 136 follow the end portions 208 of the slots 154. The forward end of the fin 112 engages the stop of the wedge 192. Thus, the fin 112 is fully deployed with a leading edge 120 facing the forward end of the missile 10 (Figure 2). The fin locking mechanism 202 cooperates with the end portions 208 of the cam slots 154 and the stop of the wedge 192 to reduce the rocking of the fin 112 relative to the cam 134 during the remainder of the missile's flight. Specifically, the wedge 192 prevents the fin 112 from coming out of the locking mechanism 202 during a forward rocking motion of the fin 112.
The deployment mechanism 114 shown in Figures 6-10 is continuously active as is the case with the deployment mechanism 14 shown in Figures 3 and 4. In other words, the deployment mechanism 114 continuously applies a force to the fins 112. This urges the fins 112 to rotate from the stowed orientation to the deployed orientation.
During the assembly of the missile, the fins 112 are assembled in or moved to the stowed orientation and placed inside a missile launch tube, for example (not shown). As a result of placing the fins 112 in the stowed orientation, the deployment mechanism 114 continuously applies a force to the pivot pin 140 along the longitudinal axis 156 of the cam 134 toward the disk 178. Without a locking mechanism to retain the fins 112 against the missile body 16 (Figure 1), the fins 112 pivot about the axially relieved portion 164 with the distal end of the fins 112 moving away from the surface of the missile 26 (Figure 1) and
engaging an inner surface of the launch tube. The inner surface of the launch tube thus prevents the fins 112 from fully deploying.
During launch, the distal ends of the fins 112 engage the inner surface of the launch tube as the missile moves down the launch tube. Once the fins 112 clear the end of the launch tube, the deployment mechanisms 114 can complete the deployment of the fins 112. The drive springs 138 urge the laterally extending end portions of cam pins 136 to move through the cam slots 154. The fins 112 pivot and then rotate with the cam pins 136 until the bases of the fins 112 engage the fin locking mechanisms 202 and the stops of the wedges 192 of the cams 134. Thus, the fins 112 fully deploy with the leading edges 120 facing the forward end of the missile 10 (Figure 1) and with a longitudinal axis 124 of each fin 112 extending substantially perpendicular to the surface of the missile 26 (Figure 2).
In an alternative embodiment, the deployment mechanism 114 may be manually or automatically activated. A retaining mechanism (not shown), such as a retaining pin, may be used to hold each fin 112 in the stowed orientation. Once the retaining pin is removed, the deployment mechanism 114 deploys the fin 112 as described in the preceding paragraph.
Figures lla-llband 12 show another assembly ofa fin 212 and another embodiment of a deployment mechanism 214. The deployment mechanism 214 is substantially the same as the previously described deployment mechanism 114 (Figure 6). However, the deployment mechanism 214 includes an alternative cam 234. In this embodiment, the disk 178 (Figure 6) in the previous embodiment is incorporated into the mounting end of the cam 234 to form a single unit. In other words, the cam 234 has a closed end 278 that performs the function of the disk 178 (Figure 6). The closed end 278 is in the shape ofa disk and has an opening 284 therethrough. The opening 284 may be shaped as two interconnecting openings with a large diameter opening 285 near an outer edge of the closed end 278 and a small diameter opening 287 near the center of the closed end 278. Surrounding the small diameter opening 287 is a recessed surface 289 for receiving the flange 248 of the pivot pin 240. The closed end 278 of the cam 234 allows the final assembly to be completed completely from the exterior. This embodiment further reduces the number of parts of the deployment mechanism 214.
The assembly, including the control fin 212 and the deployment mechanism 214 is
shown in combination with an actuator 291 in a deployed configuration in Figure 12. In this
embodiment, the cam 234 functions as an actuator shaft rotatably mounted to the actuator 291 for selectively rotating the control fin 212 about a longitudinal axis 256 of the cam 234 once the control fin 212 is in the deployed orientation. A missile guidance controller (not shown) selectively controls the actuator 291 to rotate the control fin 212 relative to the direction of airflow for controlled flight of the missile.
More specifically, as shown in Figure 12, the cam 234 is seated in the actuator 291 within an upper bearing 293 and a lower bearing 295. The cam 234 has threads on an outer surface of the lower end for receiving a threaded nut 297 thereon. The cam 234 also has an upper land or ridge 299. The upper ridge 299 engages the inner race of the upper bearing 293, and the nut 297 engages an inner race of the lower bearing 295. As the nut 297 is tightened and torqued, the two bearings 293 and 295 are trapped across a mounting block 301 of the actuator 291 and pre-loaded to secure the cam 234 to the actuator 291. This keeps the cam 234 from rattling around and allows the actuator 291 to rotate the cam 234, and thus the fin 212, at high speeds.
Now referring to Figures 13-15, yet another assembly is shown. This assembly includes a fin 312 and a deployment mechanism 314. The fin 312 has a connecting portion 350 with a spherical attachment point 351. The spherical attachment point 351 has a central notch 352, which separates the spherical attachment point 351 into two generally hemispherical portions. The spherical attachment point 351 also has a through hole 353 for receiving a cam pin 336 therein.
The spherical attachment point 351 is manufactured to fit with a very close tolerance against the inner diameter of the cam 334. This allows the spherical attachment point 351 to reduce the stress on the cam pin 336 as the fin 312 pivots and rotates from the stowed orientation to the deployed orientation. In particular, the spherical attachment point 351 reduces the stresses acting on the cam pin 336 in the fully deployed orientation of the fin 212 by transferring those stresses to the spherical attachment point 351.
At a base 372 of the fin 312, wedge shape protrusions extend from opposite faces of the fin 312 to form a key 398. The key 398 cooperates with the deployment mechanism 314 to help hold the fin 312 in the deployed orientation as will be clear from the following explanation.
The deployment mechanism 314 is substantially similar to the previously described deployment mechanism 114 (Figure 6) except as particularly described in the following paragraphs. The deployment mechanism 314 includes the cam 334, the cam pin 336, a drive spring 338, a pivot pin 340 and a disk 378 assembled as described with respect to Figures 6-10. The cam 334 has a relieved portion 364 and two laterally spaced upright sections 368 and 370. Between the laterally spaced upright sections 368 and 370 and opposite the relieved portion 364 is a keyway 355. The keyway 355 provides additional stability for the fin 312 upon full deployment and prevents or minimizes rocking of the fin 312 during the remainder of the missile's flight.
The invention thus provides a simple and reliable mechanism to both hold the fins in a stowed position and to release the fins to a deployed configuration. Further, no parts of the device are shed or broken away upon deployment of the fins, thereby minimizing or eliminating the risk of injury to the launch vehicle or operator.
Although the invention has been shown and described with respect to certain preferred embodiments, it is obvious that equivalent alterations and modifications will occur to others skilled in the art upon the reading and understanding of this specification and the annexed drawings. In particular regard to the various functions performed by the above described components (assemblies, devices, sensors, circuits, etc.), the terms (including a reference to a Ameans@) used to describe such components are intended to correspond, unless otherwise indicated, to any component which performs the specified function of the described component (i.e., that is functionally equivalent), even though not structurally equivalent to the disclosed structure which performs the function in the herein illustrated exemplary embodiments of the invention. In addition, while a particular feature of the invention may have been disclosed with respect to only one of several embodiments, such feature may be combined with one or more other features of the other embodiments as may be desired and advantageous for any given or particular application.

We claim
1. At least one aerodynamic fin (312) and a deployment mechanism (314) for a missile (10)
having the at least one aerodynamic fin (312), the deployment mechanism (314)
comprising:
a tubular cam (334) having at least one cam slot (354);
a cam pin (336) connected to the fin (312) and extending into the at least one cam slot (354); and
a spring (38/138/238/338) mountable in a missile (10) for deploying the fin from a stowed orientation to a deployed orientation that is different from the stowed orientation; and
characterized in that the fin (312) has a connecting portion (350) with a spherical attachment point (351) manufactured to fit with a very close tolerance against the inner diameter of the tubular cam (334).
2. The fin (312) and the deployment mechanism (314) as claimed in claim 1, wherein the spherical attachment point (351) has a central notch (352) which separates the spherical attachment point (351) into two generally hemispherical portions.
3. The fin (312) and the deployment mechanism (314) as claimed in claim 2, wherein the spherical attachment point (351) also has a through hole (353) for receiving the cam pin (336) therein.
4. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 3, wherein the spring (338) is an extension spring.
5. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 4, wherein the spring (338) is connected to the cam pin (336) to urge the cam pin (336) to a deployed configuration in which the fin (312) is in the deployed orientation, and
wherein the cam pin (336) is movable along and guided by the at least one cam slot (354) to pivot the fin (312) and to rotate the fin (312) from the stowed orientation to the deployed orientation.
6. The fin (312) and deployment mechanism (311) as claimed in any of claims 1 to 5, wherein the tubular cam (334) has an upper face (62) that forms a fulcrum about which the fin (312) pivots.
7. The aerodynamic fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 6, wherein the fin (312) includes a protruding key (398) adjacent an end thereof and the deployment mechanism (314) includes a keyway (355) for receiving the key to inhibit a rocking movement of the fin.
8. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 7, wherein at the base of the fin (312), wedge shape protrusions extend from opposite faces of the fin (312) to form the key (398).
9. The fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 8, wherein the cam (334) has a relieved portion (364) and two laterally spaced upright sections (368, 370) and wherein the keyway (355) is between the laterally spaced upright sections (368, 370) and opposite the relieved portion (364).
10. A missile (10) comprising the aerodynamic fin (312) and the deployment mechanism (314) as claimed in any of claims 1 to 9.
11. A missile (10) as claimed in any of claims 1 to 10, comprising cylindrical missile body (16) having an outer surface (26) and a recess (28) in the surface sized to receive the fin (312) in the stowed orientation.
12. A missile (10) as claimed in claim 10 or claim 11, comprising an actuator (291) for selectively rotating the control fin (312) about a longitudinal axis (356) once the control fin (312) is in the deployed orientation.
13. A missile (10) as claimed in claim 12, wherein the actuator (291) rotates the fin (312) relative to the direction of airflow for controlled flight of the missile.

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Patent Number

268820

Indian Patent Application Number

2491/DELNP/2004

PG Journal Number

39/2015

Publication Date

25-Sep-2015

Grant Date

17-Sep-2015

Date of Filing

26-Aug-2004

Name of Patentee

RAYTHEON COMPANY

Applicant Address

870 WINTER STREET, WALTHAM, MASSACHUSETTS 02451-1449, U.S.A

Inventors:

#	Inventor's Name	Inventor's Address
1	RUDOLPH ADOLPH EISENTRAUT	5855 SOUTH OLD SPAINSH TRAIL, TUCSON, AZ 85747, U.S.A
2	MARTIN ALLEN KEBSCHULL	3039 SOUTH NICOLETTE, AZ 85730, U.S.A
3	JOHN CHRISTOPHER PARINE	1422 SOUTH MOONFLOWER LANE, TUCSON, AZ 85748, U.S.A

PCT International Classification Number

F42B 10/14

PCT International Application Number

PCT/US2003/07931

PCT International Filing date

2003-03-14

PCT Conventions:

#	PCT Application Number	Date of Convention	Priority Country
1	10/102,032	2002-03-19	U.S.A.