Abstract:
An integrated system for missile steering, which uses both jet reaction control (JRC) and aerofin control systems, is provided with a variable coupling mechanism for adjusting the relative responsiveness of the two systems in accordance with the pressurization state of the JRC system pressure chamber. In one embodiment, the pivoting action of a joystick which actuates the gas flow control pintles of the JRC system is permitted only under sufficient pressurization of the pressure chamber. In a second embodiment, the extent to which the pintles protrude from their controllable housings is adjusted according to the pressure in the pressure chamber. In this manner, when JRC is undesirable or is unavailable, the missile aerofins are permitted their full range of motion without being constrained by the pintles.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates to rocket propelled vehicles such as missiles, and more particularly, to arrangements for steering such vehicles by a combination of steering fin control and jet reaction control. 
     2. Description of the Related Art 
     Missile control can be effected using a variety of steering schemes. One such scheme involves pivoting the thrust vectoring nozzle of the missile about a pivot point and controlling the direction of its thrust in order to provide steering in a desired direction. 
     Another method utilizes movable aerofins projecting into the airstream around the missile. This imparts to the missile the necessary forces to change its direction during flight within the earth&#39;s atmosphere and thereby effect steering control. 
     Jet reaction control (JRC) provides yet another method for steering a missile during flight, and is shown in U.S. Pat. No. 5,016,835 of Kranz. This method involves selective firing of jet nozzles disposed radially around the periphery of the missile in order to orient the missile in a desired direction. The fired jets impart an opposing reactive force on the missile and, depending on their arrangement, can serve to produce a change in direction along the yaw, pitch and/or roll axes. 
     It is also known in the art to effect missile control during flight using a combination of steering methods. One such combination, disclosed in U.S. Pat. No. 5,505,408 of Speicher et al., assigned to the same assignee as the present invention and incorporated herein by reference, relies on both jet reaction control (JRC) and control actuator fins. The two steering schemes operate in conjunction with one another to effect missile control, and yield a particularly advantageous arrangement because in some situations, when the dynamic pressure is low, such as during high attack angles or in a reduced atmosphere, the jet reaction control mechanism can compensate for the diminished effectiveness of the steerable aerofins, avoiding a compromise of missile maneuverability. 
     SUMMARY OF THE INVENTION 
     Arrangements in accordance with the present invention use an integrated missile steering system in which both jet reaction control and steerable aerofins are employed. An improved mechanical linkage between the jet reaction control and the steerable aerofins is provided, enhancing overall system performance. Use is made of a variable coupling arrangement which operates to completely decouple the two steering mechanisms or to change their relative responsiveness to steering command signals. 
     Different embodiments of the invention utilize various mechanical linkages between the steerable aerofins and the pintles which control the efflux of exhaust gases from the nozzles of the jet reaction control mechanism. These mechanical linkages can be arranged such that the ratio between the fin motion and the pintle motion can be adjusted so that small fin motions give large pintle motions. Moreover, the invention allows large pintle motions with small fin motions to be used initially in the missile flight and then, upon-burn out of the thrust vectoring gas generator, allows large fin motions without over stroking the pintle actuator. Use with a multiple burn gas generator is also contemplated, where the pintles would decouple between gas generator burns and couple during burns. 
     According to the invention, the decoupling is performed in a cost effective and highly reliable manner, allowing full motion of the aerofins without damage to the pintles or pintle drive mechanisms. Two implementations are employed, one in which a yoke plate is used, and the other in which differential area pistons in the pintles themselves are used. 
     According to the first, yoke plate arrangement, use is made of a simple mechanism which effectively unlocks the pivot bearing of the joystick lever which manipulates the individual pintles, allowing the joystick to move sideways, rather than to pivot about a point, when forces are applied thereto by the yoke plates, effectively disengaging it from the pintles. This mechanism is reliable and Low cost and is simply activated by the process of pressurizing the gas generator. Upon pressurization of the gas generator, a piston is pushed axially to capture the pivot bearing of the joystick, preventing ineffectual sideways movement of the pivot bearing and joystick and coupling the joystick to the pintles. When pressure is released at burn-out of the gas generator, forces on the joystick push the piston axially to unlatch the pivot bearing. The result is a system which is normally unlocked until the gas generator is pressurized and which stays locked during gas generator pressurization and then subsequently unlocks at depressurization. This allows full aerofin control during periods of the flight when jet reaction control is not desired or is unavailable. It also allows different ratios to be selected to optimize the response of the pintle actuators while the aerofin in motion could be restrained due to this ratio selection. An alternative embodiment uses radially, rather than axially, mounted pistons. 
     The second arrangement provides the mechanical coupling using a differential area piston in the pintle itself. This differential area piston extends the pintle to an internal hard stop when the gas generator chamber is pressurized. This allows normal functioning of the gas generator and pintle system at pressurization. Upon depressurization or burn out, the differential area piston allows the pintle to move axially when the aerofin actuator causes the pintle to contact the nozzle throat. This system is inherently simple and relies on chamber pressure to control the pintle state and allows inherent decoupling from the aerofin actuator upon depressurization. If the pintle repressurizes, the pintles are recoupled to the stick. 
     In one configuration the joystick is dispensed with and the pintle is driven by a pinion coupled directly to the actuator which operates the aerofins. A dual pintle arrangement is used, with dual differential area pistons which cause the pintles to be extended internally until they reach a hard stop. When the chamber pressure drops, the pintles are allowed to retract into the housing which supports them, thereby allowing the aerofin actuator to have larger strokes than a hard mounted pintle would. 
    
    
     DESCRIPTION OF THE DRAWINGS 
     A better understanding of the present invention may be realized from a consideration of the following detailed description, taken in conjunction with the accompanying drawings, in which: 
     FIG. 1 is a schematic perspective view, partially broken away, illustrating one particular prior art arrangement; 
     FIG. 2 is a side-sectional view of the arrangement of FIG. 1, taken along line  2 — 2  thereof and showing certain structural details; 
     FIG. 3 is a schematic view showing the mounting of a single aerofin on a missile housing; 
     FIG. 4 is a schematic cross-sectional view showing the general orientation of aerofins and yoke plates in a typical prior art arrangement; 
     FIG. 5 is a schematic side view, partially broken away, showing some of the details of the internal drive mechanism employed in arrangements such as FIG. 4; 
     FIG. 6 is a schematic side view, partially broken away, showing some of the details of the internal drive mechanism employed in a typical integrated system using jet reaction control and control actuator fins; 
     FIG. 7 is a partial view of the exterior of a missile incorporating the embodiment of FIG.  6  and depicting essentially the same portion depicted in FIG. 6; 
     FIG. 8 is an operational view of a missile incorporating the embodiments of FIGS. 6 and 7; 
     FIG. 9 is a schematic view of a joystick actuated pintle system; 
     FIG. 10A is a schematic view of the variable coupling mechanism of the invention, in the engaged position, according to a first embodiment in which a single piston is used; 
     FIG. 10B is a schematic view of the embodiment of FIG. 10A in the disengaged position; 
     FIG. 11A is a schematic view of the variable coupling mechanism of the invention, in the engaged position, according to a second embodiment of the invention in which a piston array is used; 
     FIG. 11B is a schematic view of the embodiment of FIG. 11A in the disengaged position; 
     FIG. 12A is a schematic view of the variable coupling mechanism of the invention, with the pintles in an extended position, according to a third embodiment of the invention in which a differential area piston is used; 
     FIG. 12B is a schematic view of the embodiment of FIG. 12A with the pintles in a retracted position; 
     FIG. 13A is a schematic view of the variable coupling mechanism of the invention, with the pintles in an extended position, according to a fourth embodiment of the invention: and 
     FIG. 13B is a schematic view of the embodiment of FIG. 13A with the pintles in a retracted position. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     FIGS. 1 and 2 show a prior art missile steering system in which a steerable nozzle is used to effect control of the missile. This system is known as a thrust vectoring control system (TVC). A nozzle actuation system  10  is shown in conjunction with a missile  12  having a steerable nozzle  14  mounted to a rocket motor  16  via a ball and socket joint  18 , and an encompassing skin  20  which is partially broken away to show details of the steering arrangement therein. The nozzle actuation system  10  comprises a pair of nozzle actuators  22 ,  24  which are oriented orthogonally from each other in adjacent planes which are generally transverse to the missile central axis to effect steering of the nozzle  14  relative to two orthogonal “A” and “B” axes, respectively. Thus, the actuator system  10  is able to drive the nozzle  14  about the two orthogonal axes A and B for omni-directional steering. 
     Each of the individual actuators  22 ,  24  includes a yoke plate  30  and anchoring means at opposite ends of the yoke plate for anchoring the actuator to the missile skin  20 . At one end of each yoke plate  30 , the anchoring assembly  32  comprises an anchor  34  which is affixed to the inner surface of the skin  20  and serves as a pivot mount for the yoke plate  30  via a pivot pin  36 . 
     At the opposite end of each yoke plate  30 , the anchoring arrangement comprises a gear motor  38  contained in a housing  39  which is affixed to the inner surface of the skin  20 . Projecting from the housing  39  is a shaft gear  40  which is adapted to engage the adjacent end of the yoke  30  which is fashioned with gear teeth comprising part of a sector gear  42 . 
     Completing the actuation system  10  of FIG. 1 is a yoke seat  44  which is mounted circumferentially about the nozzle  14  within the openings of the elongated yoke plates  30 . The yoke seat  44  is formed as a segment of a sphere to provide sliding contact points, such as at  46 , to support the bearing loads generated by the yoke plates  30 . The seat  44  is spherically cut and has a center on the nozzle center line at a point approximately in line with the central plane between the two yoke plates  30 . 
     Each yoke plate has an elongated central opening defined by two arms which extend about the nozzle. These arms have bearing surfaces adjacent the nozzle yoke seat for transmitting lateral forces to the nozzle  14  while permitting sliding contact with the yoke seat  44 . 
     FIG. 2 illustrates particular structural details of the nozzle system  10  of FIG. 1. A generic rocket motor is pictured having a pressure vessel volume  50  and an aft closure  52  which contains the socket for a spherical ball and socket pivot  54 . The nozzle exit cone  56  of nozzle  14  is attached to the ball portion of the pivot  54  such that the exit cone  56  is constrained to rotate with three degrees of freedom about a point  58  in the center of the ball and socket pivot  54 . 
     The spherically cut surface  60  of yoke seat  44  is threadably mounted to the outside of the nozzle  14 . The surface  60  affords a suitably strong seat for contact with the two yoke plates  30 A,  30 B at four point. Two of these points are indicated at  46 B′ and  46 B″ in FIG. 2 for the yoke plate  30 B. The yoke seat  44  is spherically cut about a point  62  located along the center line of the exit cone  56  and nominally on a plane midway between the two yoke plates  30 A,  30 B. Forces transmitted through the points of contact between the yoke plates  30 A and  30 B and the yoke seat  44  generate torque which drives rotation of the nozzle  14  about the A and B axes. 
     The A-axis actuator  22  comprises yoke plate  30 A which is attached to the missile skin structure  20  through a pivot pin  36 A. The yoke plates  30 A,  30 B are constrained to move in planes about their respective pivot pins  36  by the surrounding structure—i.e., the skin structure  20  fore and aft—as they are driven by the gear motor arrangement  38 . Each yoke plate  30 A,  30 B contains an elongated slot  64 A or  64 B. The yoke seat  44  lies within the slots  64 A,  64 B and makes contact at two points on opposite sides of each of the yoke plates  30 A,  30 B. The slots  64 A,  64 B and seat  44  are cut for a slight clearance, so that the yoke plates  30 A,  30 B are not actually in contact with the seat at both contact points at the same time, but rather will contact one point or the other depending upon the direction of applied forces. Each yoke plate  30 A,  30 B has gear teeth  70 A or  70 B cut into the plate at one end to establish a sector gear portion which is driven by a cluster shaft pinion  72  (FIG.  1 ). The cluster shaft is mounted by bearings  74 ,  76  to the missile skin structure  20 . The A-axis drive motor  80 A is mounted on tabs  82 A of the missile skin structure  20 . The motor shaft pinion  84 A drives the cluster shaft  40 A. Clearance slots are cut into the yoke plates  30 A,  30 B to allow long rotation of the yoke plates without interference from the other axes cluster pinions  72 . 
     The B-axis drive is essentially identical to the A-axis drive. The B-axis yoke plate  30 B is positioned next to, but in front of, the A-axis yoke plate  30 A. Its pivot pin  36 B is similarly attached to the missile structure  20 , and yoke plate  30 B has sector gear teeth  70 B driven by an engaged pinion  72 B on shaft  73 B. 
     Rather than pivoting each of the yoke plates at one of its anchoring points, a mounting arrangement in which the yoke plates are permitted to translate along orthogonal axes can be provided (FIG.  4 ). Additionally, in combination with the thrust vectoring control (TVC) system using a pivotable nozzle, steerable aerofins can be employed to augment missile steering control in an integrated steering arrangement, illustrated in prior art FIGS. 3-5. FIG. 3 is a schematic diagram representing a missile  110  with an aerofin assembly  112  installed thereon. The assembly  112  comprises an aerofin  120  pivotably installed on a base plate  114  which is secured to the skin  116  of the missile  110  by means of mounting bolts  118 . The aerofin  120  is affixed to an internal drive mechanism by mounting bolts  122 . The exhaust nozzle of the missile  110  is represented schematically at  124 . The pivotable mounting of the nozzle  124  corresponds to that which is shown in FIGS. 1 and 2. 
     FIG. 4 is a schematic diagram illustrating the drive elements of the steering control system of the prior art. A pair of orthogonally oriented yoke plates  130 ,  132  are shown bearing against the steerable nozzle  124  to control thrust direction in a manner similar to that of the prior art arrangement depicted in FIGS. 1 and 2. A principal difference from that device is that each of the yoke plates  130 ,  132  is free to move in response to rotational forces applied at both opposite ends thereof, rather than being pivotally anchored at one end as indicated in FIG.  1 . 
     The details of the yoke plate drive assemblies are shown for the unit A at the position of the aerofin assembly  112 . A rack and pinion gear assembly  136  comprises a curved rack gear  138  on a rack carrier  140 . The carrier  140  is curved on its outer surface to match the curvature of the missile shell  142  and is adapted to slide circumferentially relative to the missile shell  142  as it is driven by the spur or pinion gear  144 . The corresponding end of the yoke plate  132  is provided with a U-shaped recess  146  in which the rack carrier  140  is mounted, bearing against side walls  148  of the recess  146 . This arrangement is repeated at the other three aerofin stations B, C and D located at 90 degree spacings about the missile. 
     In FIG. 4, the broken line outline  150  indicates the typical launcher envelope for such a system. It will be apparent that, as the pinion gear  144  is driven to rotate, it moves the rack carrier  140  either clockwise or counterclockwise, depending upon the direction of rotation of the pinion gear  144 . Corresponding movement of the yoke plate  132  moves the nozzle  124  off axis, thereby changing the direction of the thrust to effect steering of the missile. 
     FIG. 5 illustrates schematically the details of the combination drive arrangement for an aerofin in  112  and a yoke plate  132 . This view shows the combined aerofin and TVC dual pinion gear  160  having a central drive gear  162  mounted on a common shaft with pinion gear  144  and a bevel pinion gear  164 . The shaft of the dual pinion gear  160  is mounted in bearings  166 . 
     A bevel gear  170  is directly connected to the aerofin  120  and is coupled to the bevel pinion gear  164 . Gear  170  is mounted for rotation in upper and lower bearing  172 ,  174 . An electric motor  180  has an output shaft coupled to drive the gear  162  which in turn produces rotation of both the bevel gear  170  and the pinion gear  144 , thus driving both the aerofin  120  and the rack  140 . This in turn drives the yoke plate  132 . A feedback transducer  182  is connected to the aerofin bevel gear  170  by a shaft  184 , thereby providing aerofin position data for the control system of the drive arrangement  100 . The coupling between the motor  180  and the gear  162  is represented by the block  178 . This preferably incorporates a speed reducing gear train to transform the motor&#39;s relatively high speed and low torque into low speed and high torque. Such speed reducers are known in the art; details are omitted from FIG. 5 for simplicity. 
     A different integrated steering arrangement, which uses, a combination of aerofin in and jet reaction control (JRC), is represented schematically in FIGS. 6-8. FIG. 6 shows an actuator assembly like that depicted in FIG. 5, except that here the actuator assembly serves to control an associated auxiliary jet steering system rather than the thrust vector control system of the main nozzle as previously described. 
     The actuator assembly portion of FIG. 6 to the left of the broken line A—A is the same as that shown in FIG.  5  and the same reference numerals are used to identify like elements. It should be clear, of course, that there are four of the assemblies like the one depicted at the bottom of FIG. 6, one for each of four fins  120  mounted at  90  degree angles about the missile  110 . 
     The jet reaction control portion of the arrangement of FIG. 6 is shown comprising a JRC housing  200  mounted just aft of the yoke plates  206 ,  208  which are positioned to control the movement of the valve control puck  204 . These elements correspond to or are equivalent in operation to the yoke plates  130 ,  132  and the steerable nozzle  124  in the FIG. 4 representation of the first preferred embodiment, described hereinabove. 
     The housing  200  encompasses four rocket nozzles  202  and four associated rocket valves  210  situated about a central pressure inlet  216  from a rocket motor or other pressure source  220 . These rocket nozzles and valves may be oriented to exhaust directly behind the aerofins  120 , as indicated in FIG. 6 or they may be angularly displaced therefrom as desired, for example, displaced by 45 degrees so that the nozzles exhaust midway between the aerofins  20 . 
     Each valve  210  is generally cylindrical with a bullet nose  214  bearing against a valve seat  215 . The valve  210  is hollow and contains a spring  218  therein for urging nose  214  of the valve  210  against the seat  215  to close off the associated passage from the pressure inlet  216  to a corresponding nozzle  202 . To one side of the valve  210  is a valve arm  212  which bears against the outer surface of the valve control puck  204 . Thus as the puck  204  is moved off the central axis of the missile by the actuator assembly, as previously described, it drives one or another of the valve arms  212  and associated valve  210  radially outward, thereby drawing the nose  214  away from the seat  215  to a valve-open position, as indicated in the broken line of the lower valve in FIG. 6, so that gas from the pressure inlet  216  connects through that valve passage to the bottom nozzle  202  in FIG.  6 . 
     The effect of opening one of the valves  210  in this manner is illustrated in FIGS. 7 and 8. FIG. 8 shows a portion of a missile body with aerofins  120  and a nozzle  202  mounted directly behind the aerofins. The operation of this system is represented at FIG. 8 where the portion of FIG. 7 is shown installed on the missile as a canard system. The aerofins  120  are shown rotated to cause a force pushing the nose of the missile  110  down. Similarly, the exhaust  203  from the nozzle  202  in the uppermost position operates to produce the same effect, driving the nose of the missile downward to produce a directional change indicated by the arrow A. 
     In accordance with a first embodiment of the invention, illustrated in FIGS.  9 — 13 , the valves  210 , hereinafter referred to as pintles and designated by reference numeral  230 , are actuated by means of a pivotably mounted joystick  232  rather than by control puck  204 . Joystick  232 , having an optional flexible seal  250 , is movably mounted for engagement with pintles  230  disposed radially therearound. Joystick  232  is actuated by yoke plates  234 ,  236  of an actuator assembly similar to that described above. The pivoting motion of the joystick  232  can be selectively coupled to the pintles  230  by controlling its pivoting action at pivot bearing  233 . In this manner, selective control of the flow of exhaust gases from pressure chamber  238  through nozzles  240 , in response to movement of the yoke plates and in coordination with the aerofins  120 , is attained. 
     FIG. 10A shows the variable coupling mechanism of the invention in the engaged position. A pivot seat  246  having bearing surface  248  is mounted on a translating piston  244 . Piston  244  is mounted in piston bearing  242  and translates axially therein. Piston bearing  242  is in communication with pressure chamber  238 , permitting the axial position of the pivot seat  246  and the piston  244  to change according to pressure in pressure chamber  238 . Under pressurization conditions, pivot seat  246  is forced into the engaged position of FIG. 10A to thereby contact pivot bearing  233  and provide a pivoting surface for the pivot bearing  233 , limiting the motion of joystick  232  to a pivoting action. In this configuration, the motion of aerofins  120  via yoke plates  234 ,  236  is effectively coupled to thrust nozzles  240 , with movement of the aerofins causing corresponding thrusting of the jet reaction control system to achieve integrated steering of the missile. 
     When pressure chamber  238  depressurizes, piston  244  and pivot seat  246  are caused to translate axially away from pivot bearing  233 , by forces on the joystick  232 , to the position illustrated in FIG.  10 B. This disables the pivoting action of joystick  232 , decoupling the motion of yoke plates  234  and  236  from pintles  230 . 
     In a second embodiment of the invention depicted in FIGS. 11A and 11B, rather than a single piston  244 , a pivot seat array  253  is used to provide the pivoting surface for pivot bearing  233  and limit the motion of joystick  232 . The pivot seat array  253  is mounted on a piston array  252  and translates in array bearing  254 , which is in communication with pressure chamber  238 . Pivot seat array  253  and piston array  252  operate to couple and decouple the motion of yoke plates  234 ,  236  from pintles  230  in accordance with the pressurization state of the pressure chamber. FIG. 11A depicts the pivot seat array  253  in the engaged position, while FIG. 11B depicts the array in the disengaged position. 
     A third embodiment of the invention encases pintles  230  within translating pintle housings  256  to form differential area pistons. Pintle housings  256  are actuated by joystick  232  and translate along housing bearings  258  to control the exhaust stream through nozzles  240 . Contained within each pintle housing  256  is expansible subchamber  260  which has as a boundary thereof one edge of pintle  230 . Subchamber  260  communicates with pressure chamber  238  via channels  262  formed in pintles  230 . When pressure chamber  238  is pressurized, pressure in subchamber  260  forces pintle  230  outward a corresponding distance, allowing a normal response of the pintles to yoke plates  234  and  236  and joystick  232 . In this configuration, depicted in FIG. 12A, small motions of the yoke plates and joystick are sufficient to provide gas flow control through nozzles  240  and effect missile steering. 
     Upon depressurization or burn out, the differential area piston allows the pintle  230  to retract into pintle housing  256  when the joystick  232  presses pintle  230  against nozzle throat  266 . In this manner, the arrangement decouples the jet reaction control (JRC) system from the aerofin control during periods of depressurization. The decoupling permits greater range of motion of the aerofins as they are no longer inhibited by the limited range of motion of the pintles  230  to which the aerofins were coupled. Moreover, the system permits recoupling when the pressure chamber  238  repressurizes in situations where the need for extreme aerofin motions is reduced and jet reaction control is desired. The decoupled configuration is illustrated in FIG.  12 B. 
     In an alternative embodiment shown in FIGS. 13A and 13B, pinion gears  268  replace joystick  232 . Two pinion gears  268 , each associated with a pair of pintles  230  mounted in a housing  272 , couple the aerofin control system to the jet reaction control system. The housings  272  are each provided with a rack gear  270  for engagement with the pinion gears  268 . A subchamber  260  is formed in each housing and optionally contains a bulkhead  274  therein. The subchamber is bounded at two opposing ends by pintles  230 , which pintles have channels  262  formed therein to permit communication of the subchambers  260  with the pressure chamber  238 . Pressure in pressure chamber  238  causes outward extension of pintles  230  along pintle bearings  264  formed in the housings  272 , allowing normal control of the gas flow through nozzles  240  by the pintles in response to actuation of housings  272  by pinion gears  268 . 
     Under depressurization conditions, depicted in FIG. 13B, pintles  230  are permitted to retract into the housings  272  when pressed against the nozzle throats  266 , reducing the response of the jet reaction control system to pinion gears  268 . This configuration affords maximum movement and control of the aerofins by removing constraints imposed by the otherwise limited motion of the pintles  230 . The arrangement thus achieves a simple, variable response system which adjusts to the exigencies of the particular missile flight conditions. 
     Although there have been described hereinabove various specific arrangements of a Variable Coupling Arrangement for an Integrated Missile Steering System in accordance with the invention for the purpose of illustrating the manner in which the invention may be used to advantage, it will be appreciated that the invention is not limited thereto. Accordingly, any and all modifications, variations or equivalent arrangements which may occur to those skilled in the art should be considered to be within the scope of the invention as defined in the annexed claims.