Abstract:
Disclosed is a thrust termination device for a solid rocket motor, which terminates the net thrust by the reverse thrust of the rocket motor produced from the emission of the combustion gas in the reverse direction, when the stage separation signal is transferred at the normal thrust state of the rocket motor. An object of the present invention is to provide a thrust termination device for a rocket motor, which can contrive to accomplish the structural safety and mechanical sealing performance at the combustion chamber condition of the high temperature and high pressure, and easily remove the thrust termination device even at the low pressure state and open the trust termination ports successively with very small impacts when the thrust termination is commanded.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     1. Field of the Invention  
         [0002]     The present invention relates to a thrust termination device of a solid rocket motor, which terminates the net thrust by the reverse thrust of the rocket motor produced from the emission of the combustion gas through the opened thrust termination port in the reverse direction, when the stage separation signal is transferred at the normal thrust state of the rocket motor, in particular, to a thrust termination device for a rocket motor, which can contrive to accomplish the structural safety and the mechanical sealing performance at the combustion chamber condition of the high temperature and high pressure, and easily remove the thrust termination device even at the low pressure state and open the trust termination ports successively with very small impacts when the thrust termination is commanded.  
         [0003]     2, Background of the Related Art  
         [0004]     In general, the solid rocket motor comprises a combustion chamber for combusting solid propellant, and a nozzle for the exit of exhaust gases which makes acceleration of rocket in the opposite direction to the gases flow direction. And the thrust, which is the product of acceleration and mass of rocket, have same direction of acceleration of rocket motor.  
         [0005]     There have been developed several kinds of rocket motors, and the principle and mechanism of propulsion are different for each rocket motor. The most widely used rocket motor among them is a chemical rocket motor with an energy source of chemical reaction (combustion) of solid propellant or liquid propellant.  
         [0006]     For instance, in a rocket motor using the solid propellant as a fuel, ignition of the solid propellant is performed by a igniting device, and gases produced from the combustion of the solid propellant are exhausted through the outlet of the nozzle, and the retroaction against exhausted gases becomes the thrust of the rocket motor in the normal propulsion state till the practice of the command for the thrust termination. Thus, the rocket motor is provided with a thrust termination device for producing thrust opposing to the thrust of the nozzle so as to neutralize the normal thrust of the nozzle.  
         [0007]     When the rocket motor provided with the thrust termination device as described above is schematically reviewed, it comprises a combustion tube for a combustion chamber with a solid propellant charged; a nozzle mounted at the rear of the combustion tube; a dome portion of the combustion tube formed at the front side of the combustion tube; an igniting device mounted at the center of the dome portion of the combustion tube toward the combustion chamber; and a plurality of thrust termination devices mounted at the dome portion of the combustion tube.  
         [0008]     According to the rocket motor, normal thrust can be produced at the nozzle at the normal state, however, in this instance, a thrust termination port formed at the dome portion of the combustion tube to communicate with the combustion chamber is maintained to be closed by the thrust termination device. Further, when the command of thrust termination is practiced, the thrust termination port closed by the thrust termination device is instantly opened to produce thrust greater than the normal thrust of the nozzle through the thrust termination ports to thereby impede the forward movement of the rocket motor and reduce the pressure in the combustion tube, so that forward thrust cannot be produced any more in the rocket motor. In this instance, the sum of the sectional areas of the opened thrust termination ports is greater than that of the sectional area of the nozzle throat to thereby make the thrust produced from the thrust termination port greater than the normal thrust of the nozzle, so that the movement of the rocket motor in the forward direction can be impeded.  
         [0009]     Conventionally, a Pyrotechnic devices (explosion bolt) described in U.S. Pat. No. 5,400,713 are mostly used in the thrust termination device.  
         [0010]     In this regard, it is required for the thrust termination device to be sufficiently safe in structure before the operation, operated rapidly at the operation and reproduced completely. However, when the thrust termination device such as the pyro device was operated, there were produced substantially big impacts, and fragments of the operated device were dispersed, so that the missile had to be affected from the substantial impact during the operation of the thrust termination device to thereby affect badly to the flying stability of the missile and the precision of the control after the thrust termination.  
         [0011]     Accordingly, it is very important to reduce the magnitude of impacts at the time of the operation of the thrust termination device.  
       SUMMARY OF THE INVENTION  
       [0012]     Therefore, the present invention has been made to decrease the magnitude of impacts occurring in the prior arts, and an object of the present invention is to provide a thrust termination device for a rocket motor, which can contrive to accomplish the structural safety and the mechanical sealing performance at the combustion chamber condition of the high temperature and high pressure, and easily remove the thrust termination device even at the low pressure state and open the trust termination ports successively with small impacts and smoothly when the command of thrust termination is performed.  
         [0013]     To accomplish the above objects, according to the present invention, there is provided a thrust termination device for a rocket motor, which is constructed to close a plurality of thrust termination ports formed at a dome portion of a combustion tube mounted at the front of the rocket motor to communicate with a combustion chamber of the rocket motor in a sealing structure at a normal thrust state, and to open them respectively when a command of thrust termination is performed, the thrust termination device is characterized by comprising a main cylinder opened at one end and supported by the thrust termination ports to close them, and formed with a vent opening at a bottom thereof; a primary piston inserted into the vent opening from the inside of the main cylinder, and transmitted with the pressure in the combustion chamber; a secondary piston mounted around the primary piston in the inside of the main cylinder, and transmitted with the pressure in the combustion chamber through the vent opening by the movement of the primary piston due to the pressure in the combustion chamber; a pressure load transmitting device mounted around a front end of the secondary piston and engaged with the secondary piston; a restriction pin for penetrating through the pressure load transmitting device, and the primary/secondary pistons to restrict them, and having a drawing wire connected at one end thereof and pulled to relieve the restriction state of them, when the command of thrust termination is performed; and a cross plate mounted at the front of the pressure load transmitting device and the main cylinder, and engaged with an inner peripheral surface of the thrust termination port at the edge thereof to restrict the pressure load transmitting device and the main cylinder.  
         [0014]     According to the present invention, the thrust termination port is formed at the inner peripheral surface thereof with four spline grooves for inserting the edges of the blades of the cross plate and four sills for supporting the edges of the blades according to the rotation angle of the cross plate.  
         [0015]     Further, the restriction pin is a split pin formed to be separated from the pressure load transmitting device, and the primary/secondary pistons by the mechanical drawing force of the drawing wire. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0016]     The above and other objects, features and advantages of the present invention will be apparent from the following detailed description of the preferred embodiments of the invention in conjunction with the accompanying drawings, in which:  
         [0017]      FIG. 1  is a cross-sectional view for showing an example of a rocket motor applied of a thrust termination device of the present invention.  
         [0018]      FIG. 2  is an enlarged view of a principal portion in  FIG. 1 .  
         [0019]      FIG. 3  is a partial front view for showing a rocket motor applied of the thrust termination device according to the present invention. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
       [0020]     Reference will now be made in detail to the preferred embodiment of the present invention with reference to the attached drawings. Prior to the detailed description of the present invention, it should be confirmed that the terms or words used in the specification and claims of the present invention are construed as meanings and concepts conforming to the technical spirit of the present invention on the basis of a principle that the inventors can define the concept of the term properly for explain their invention with the best method.  
         [0021]     In  FIG. 1 , an example of a rocket motor applied of a thrust termination device of the present invention is shown.  
         [0022]     As shown in the drawing, the rocket motor applied of the thrust termination device  200  of the present invention, comprises a combustion tube  100  forming a combustion chamber  112  and charged with a solid propellant therein, a nozzle  120  mounted at the rear of the combustion tube  100 , a dome portion  130  of the combustion tube mounted at the front of the combustion tube  100 , an ignition device  140  mounted at the center of the dome portion  130  of the combustion tube toward the combustion chamber  112 , and a plurality of thrust termination devices  200  mounted at the dome portion  130  of the combustion tube. For instance, an example is shown that two thrust termination devices  200 ,  200  are mounted in the dome portion  130  of the combustion tube in a symmetrical structure with respect to the center axis C of the rocket motor.  
         [0023]     The thrust termination device  200  according to the present invention is constructed of a mechanical structure that the thrust termination ports  132  formed at the dome portion  130  of the combustion tube  100  to communicate with the combustion chamber  112  are respectively closed in a sealing structure at the normal thrust state of the nozzle  120 , and then respectively opened to counterbalance the normal thrust by means of the reverse thrust, when the command of thrust termination is performed. The thrust termination ports  132  can be formed at the dome portion  130  of the combustion tube at 0° and 180° positions, if two ports are formed. Further, it is preferable that the thrust termination ports  132  are formed to have inclination angles of +45° and −45° with respect to the center axis C of the rocket motor, and is communicated with the combustion chamber  112  through the respective passages  114 . Also, in order to counterbalance the normal thrust of the nozzle  120  by means of the reverse thrust, the sum of the sectional areas of the thrust termination ports  132  is preferable to be greater than the sectional area of the throat  122  of the nozzle  120 .  
         [0024]     As shown in  FIGS. 1 through 3 , the thrust termination device  200  for the rocket motor of the present invention, comprises a main cylinder  210 , a primary/secondary pistons  220 ,  230 , pressure load transmitting device  250 , a restriction pin  260 , and a cross plate  270 .  
         [0025]     As shown in  FIG. 2 , the main cylinder  210  is supported by the thrust termination device  132  to close it, constructed to be opened to the outside (in other words, outer axial direction of the thrust termination device  132 ) at one end, and is formed with a vent opening  212  at the bottom center. The main cylinder  210  can be supported by the thrust termination port  132  in a sealing structure by means of an O-ring  280  at the outer peripheral surface.  
         [0026]     Further, the primary piston  220  is constructed that one end of it is inserted into the vent opening  212  from the inside of the main cylinder  210 , and the outer peripheral surface of a portion inserted into the vent opening  212  is supported by the main cylinder  210  in a sealing structure. A stopper  222  is the radially protruded portion at the outer peripheral surface of the primary piston  220 . In this regard, the stopper  222  acts to locate the primary piston  220  in the proper position into the vent opening  212 , and, as will be described hereinafter, to transmit the pressure to the pressure load transmitting device  250  at the time of the movement of the primary piston  220 .  
         [0027]     Also, the secondary piston  230  is mounted around the primary piston  220  from the inside of the main cylinder  210 , to thereby be supported by the main cylinder  210  and the primary piston  220 . In other words, the inner peripheral surface of the secondary piston  230  can be supported against the outer peripheral surface of the primary piston  220  by means of the O-ring  280 , and the outer peripheral surface of the secondary piston  230  can be supported against the inner peripheral surface of the main cylinder  210 . In this regard, the O-rings  280  concurrently function to prevent the gas leaking through the thrust termination port  132  from the combustion chamber  112  at the normal thrust state. In addition, the secondary piston  230  has a boss  232  at the center, into which the primary piston  220  is inserted. The secondary piston  230  can successively move in the moving direction of the primary piston  220  by the gas pressure in the combustion chamber  112  transmitted via the vent opening  212 , when the primary piston  220  moves toward the inside of the main cylinder  210  from the vent opening  212  by the gas pressure in the combustion chamber  112 .  
         [0028]     Further, the pressure load transmitting device  250  is mounted around the front end of the secondary piston  230 , that is, around its boss  232 , to be engaged with the secondary piston  230 . The relative movement of the pressure load transmitting device  250  and the secondary piston  230  is restricted by means of two securing screws  240  radially assembled in opposite positions at both sides of the pressure load transmitting device  250  and the secondary piston  230 . Accordingly, the pressure load transmitting device  250  can move together with the secondary piston  230 . Also, the pressure load transmitting device  250  can be engaged with the stopper  222  of the primary piston  220 , so that more ejecting force can be transmitted from the stopper.  
         [0029]     The restriction pin  260  penetrates the pressure load transmitting device  250 , the secondary piston  230 , and the primary piston  220  in about the radial direction to thereby restrict them. In this regard, the drawing wire  262  is connected to one end of the restriction pin to relieve the restriction state when the thrust termination command is performed. A split pin is used for the restriction pin  260  to separate the pressure load transmitting device  250 , and the primary/secondary pistons  220 ,  230  by such drawing. That is, the protruded ends of the restriction pin  260  are bent after the insertion through the pressure load transmitting device  250 , and the primary/secondary pistons  220 ,  230 , and the drawing wire  262  is bound to a head of the restriction pin  260 .  
         [0030]     Also, the cross plate  270  is mounted at the front of the pressure load transmitting device  250 , and supports the pressure load transmitting device  250  and the main cylinder  210  thereof by the engagement of the edge of the cross plate with the inner peripheral surface of the thrust termination port  132 . The cross plate  270  should be fabricated with very high strength structure, and also it should be easily separated even at very low inner pressure load transmitted from the combustion chamber  112  via the pressure load transmitting device  250 , after the removal of the restriction pin  260  at the time of the perform of the command of the thrust termination to thereby open the thrust termination port  132 .  
         [0031]     As shown in  FIG. 2  and  FIG. 3 , four spline grooves  134  receiving the edges of the respective blade of the cross plate  270 , and four sills  136  for supporting the edges of the blades are formed at the inner peripheral surface of the thrust termination port  132  depending on the rotation angle (that is, 45° rotation) of the cross plate  270 . Accordingly, when the cross plate  270  is inserted into the thrust termination port  132  so that edges of the blades can be inserted into the spline grooves  134 , if the cross plate  270  is rotated by 45°, the edges of the blades of the cross plate  270  are located inside the sills  136  so that the thrust termination device  200  of the present invention cannot be separated to the outside of the thrust termination port  132 .  
         [0032]     Next, the action of the thrust termination device for the rocket motor according to the present invention as constructed above will be explained below.  
         [0033]     When the thrust termination device  200  for the rocket motor according to the present invention is to be mounted, at first, the main cylinder  210  is inserted into the thrust termination port  132 , then, the primary/secondary pistons  220 ,  230  and the pressure load transmitting device  250  are assembled to the main cylinder  210 , and the restriction pin  260  bound of the drawing wire  262  is inserted into the primary/secondary pistons  220 ,  230  and the pressure load transmitting device  250 , and the protruded ends of the restriction pin  260  are bent toward the outer peripheral surface of the pressure load transmitting device  250 , to thereby prevent the falling out of the restriction pin  260 . Further, after the insertion of the respective blade of the cross plate  270  into the spline groove  134 , the cross plate  270  is rotated so that the blades of the cross plate  270  are engaged with the sills  136  to thereby complete the assembly of the thrust termination device  200  of the present invention.  
         [0034]     Thus, after the thrust termination device  200  of the present invention has been mounted to the rocket motor, if the solid propellant is ignited by means of the ignition device  140 , gases in the combustion chamber  112  is exhausted through the nozzle  120  to produce the thrust. At the normal thrust state of the nozzle  120 , the cross plate  270  is subjected to the pressure load transmitted in the thrust termination port  132 , and restricts the outward movement of the pressure load transmitting device  250  and the main cylinder  210 .  
         [0035]     At such normal thrust state of the nozzle  120 , when the command of the thrust termination is performed, the drawing wire  262  is pulled with the mechanical structure, then the bent both ends of the restriction pin  260  of the split pin structure are spread to separate the restriction pin  260  from the pressure load transmitting device  250  and the primary/secondary pistons  220 ,  230 .  
         [0036]     As described above, when the restriction pin  260  is separated, the primary piston  220  becomes to move from the vent opening  212  to the inside of the main cylinder  210  by the pressure of the combustion gas in the combustion chamber  112 , and the pressure of the combustion gas in the combustion chamber  112  is transmitted to the lower side of the secondary piston  230  through the vent opening  212 , so that the secondary piston  230  can also move in the direction identical with that of the primary piston  220 . Accordingly, since the pressure load transmitting device  250 , firstly transmitted with the pressure load through the stopper  222  of the primary piston  220 , is engaged with the secondary piston  230 , the pressure load transmitting device  250  is pushed toward the cross plate  270 . Therefore, the cross plate  270  is deformed because the pressure load of the combustion gas is transmitted to the center of the cross plate  270  through the pressure load transmitting device  250 , and the blades of the cross plate are separated from the sills  136  so that the main cylinder  210 , the primary/secondary pistons  220 ,  230 , and the pressure load transmitting device  250  are all pushed to thereby open the thrust termination port  132 .  
         [0037]     Accordingly, the net thrust can be terminated because reverse thrust can be produced in the direction opposing to the normal thrust through the nozzle  120 . That is, when the thrust termination ports  132  are opened as described above, the sum of the sectional areas of the thrust termination ports  132  becomes to be greater than the sectional area of the throat of the nozzle  120  to thereby produce the reverse thrust surpassing the thrust produced at the nozzle  120 , and the lower motor (not shown) is separated, and the pressure of the combustion chamber  112  is reduced, and in some cases combustion is terminated.  
         [0038]     According to the thrust termination device for the rocket motor of the present invention constructed as described above, it is possible to accomplish the best performance in the flying stability and the control precision at stage separation of a missile or thrust termination of the rocket motor, by the smooth and stable opening of the thrust termination port  132  without big impact through the small drawing force for removing the restriction pin  260 , and the progressive operation of the parts.  
         [0039]     Further, it is possible to increase the structural safety of the thrust termination device and perform the smooth opening of the thrust termination port  132  by the progressive actions using the multi-piston structure.  
         [0040]     While the present invention has been described with reference to the particular illustrative embodiments, it is not to be restricted by the embodiments but only by the appended claims. It is to be appreciated that those skilled in the art can change or modify the embodiments without departing from the scope and spirit of the present invention.