JP4179157B2 - Two-stage thrust rocket motor - Google Patents

Description

  The present invention relates to a two-stage thrust type rocket motor that suppresses the structural weight, thereby extending the range and preventing the generation of the second-stage thrust from being delayed.

In recent years, a two-stage thrust type rocket motor has been proposed from the viewpoint of improving the performance of the rocket motor such as extending the range of the flying object.
This is because two propellants are loaded into the combustion chamber of a rocket motor having one injection nozzle, and after the first propellant is activated, the second propellant is activated after an arbitrary time. Two thrusts can be generated. This requires the two propellants to be isolated from each other so that the first propellant is burning and the other propellant is not ignited.
As a specific example of a rocket motor developed for such a purpose, for example, a rocket motor as shown in FIGS. 9 and 10 is known (see Patent Document 1).

  The conventional two-stage thrust rocket motor shown in FIG. 9 has a first propellant 34 on the rear side of a combustion chamber 33 provided with Porlabs 31 and 32 for attaching an igniter assembly and an injection nozzle 30 to the front end and the rear end, respectively. The second propellant 35 is loaded on the front side, and the combustion chamber 33 and Porlabus 31 are forward insulated so that the second propellant 35 does not ignite within a predetermined time interval after the first propellant 34 starts to burn. The body 36 is bonded, and this insulator is also bonded to the inner surface of the hollow second propellant 35. Further, a diaphragm 37 bonded to both propellants is disposed between the first propellant 34 and the second propellant 35. The front insulator 36 and the diaphragm 37 are set so that the second propellant 35 does not ignite a predetermined time before the first propellant 34 starts to burn. The second propellant 35 has a propellant support 38 made of a foam material disposed in the center of the front end. The injection nozzle 30 is attached to Porlabus 32.

FIG. 10 is an enlarged view of a mounting portion of the igniter assembly used in this example.
The igniter assembly is composed of an igniter block 39 and an igniter case 40 protruding from the block, and is attached to a porlabus 31 at the front end of the combustion chamber.
The first igniting agent 41 for igniting the first propellant is in the igniter case 40, and the second igniting agent 42 for igniting the second propellant is in the annular chamber 43 formed from the igniter block 39 and the wall surface of Porlabus 31. Porlabus 31 has a nozzle port 51 that allows the ignition gas of the second igniter 42 to flow into the combustion chamber. Thus, after the first propellant starts to burn by the ignition gas of the first igniting agent 41, the second propellant is not ignited by the front insulator and the diaphragm before a predetermined time.

  After a predetermined time elapses, the second igniter 42 ignites, and the ignited gas collides with the propellant support 38 from the annular chamber 43 through the nozzle port 51 and melts the propellant support. During this process, the ignition gas ignites the second propellant 35, destroys the diaphragm, and generates a second thrust. In this way, the two-stage thrust rocket motor of FIG. 9 can generate two independent thrusts.

  However, in the two-stage thrust rocket motor as shown in FIG. 9, the structure of the front end portion of the combustion chamber becomes large due to the formation of the annular chamber 43, and the structure weight increases. In addition, since combustion gas flows into the annular chamber 43 through the nozzle port 51 during the second propellant combustion, the heat insulating material installed to ensure the heat insulation of the inner surface of the chamber becomes thick, which also causes the front end of the combustion chamber. The structural weight increases as the partial structure increases. Due to these increases in structural weight, the flying object could not obtain a sufficient acceleration and had problems such as a short range. Further, since the second propellant 35 ignited by the ignition gas from the annular chamber 43 is limited to a narrow area near the nozzle port 51, the second propellant 35 is bonded to the front insulator 36 or the diaphragm 37. When the end face is a combustion face, it takes time for the entire combustion face to ignite, and as a result, flight performance problems such as speed reduction due to the delay in the generation of the second stage thrust have also occurred.

Japanese Patent No. 3321778 (pages 1 to 3, FIGS. 2 and 3)

  An object of the present invention is to provide a two-stage thrust type rocket motor that suppresses the structural weight, extends the range, and prevents the generation of the second-stage thrust from being delayed.

  As a result of intensive investigations in view of the problems in the previous period, the present inventors have determined that the second ignition housing 13 having one or more through holes in the wall surface at a position where the hollow wall surface of the second propellant 7 is directly ignited. 2 It has been found that the problem can be solved by disposing in the hollow portion of the propellant, and the present invention has been completed.

In the first invention, the injection nozzle 4 is arranged at the rear, and the first combustion chamber 1 and the second combustion chamber 2 are connected in series in this order directly or via the joint 3 from the injection nozzle 4 to the front. The first propellant 6 having a hollow portion penetrating back and forth in the first combustion chamber and the second propellant 7 having a hollow portion penetrating front and rear in the second combustion chamber are disposed respectively. A first blockade in which one or more through holes are provided in a wall surface at a position where the propellant 6 is selectively ignited, and a mechanism for igniting the first igniter 10 is attached to the front end of the first ignition housing 12. A first ignition housing 12 having a body 14 and loaded with the first igniting agent 10 is disposed on the front and rear central axis in the first combustion chamber, and the hollow portion wall surface of the second propellant 7 1st or more through-holes in a wall surface in the position to ignite directly, and 2nd ignition to the front side edge part of the housing | casing 13 for 2nd ignition The second ignition housing 13 having a second sealing body 15 attached with a mechanism for igniting 11 and having the second ignition powder 11 loaded therein is disposed in the hollow portion of the second propellant 7. The diaphragm 9 with the outer peripheral end fixed at the entire circumference of the rocket motor body and the inner peripheral end fixed between the first ignition housing 12 and the second ignition housing 13 is the first propulsion. The second propellant is disposed so as to isolate the medicine 6 from the second propellant 7, and the diaphragm 9 is not broken or lost by the generated gas of the first igniter 10 and the combustion gas of the first propellant 6. In the two-stage thrust rocket motor that breaks or disappears with the combustion gas 7 and the gas generated by the second igniting agent 11 , the mechanism for igniting the first igniting agent 10 includes the first initiator 26, the second sealing body 15, and the second The first through the ignition housing 13 and the first sealing body 14 connected to it. And stretching out to the inside the ignition housing 12, and a two-stage thrust rocket motor, characterized in that it consists of tube 23 having a JoSoyaku 24 therein for propagating ignition of the initiator 26.

The second invention is characterized in that the diaphragm 9 is EPDM rubber, silicone rubber, EPDM rubber or silicone rubber containing organic fibers, or EPDM rubber or silicone rubber containing inorganic fibers. It is a stage thrust type rocket motor.

A third invention is any Kano two-stage thrust of the first to the second invention, characterized in that the diaphragm 9 is bonded in the hollow wall surface of the rear end surface or the second propellant 7 of the second propellant 7 Type rocket motor.

The fourth invention is characterized in that a mechanism for igniting the second igniting agent 11 includes an initiator 22 and a second auxiliary agent 21 in the second sealing body 15 that propagates ignition of the initiator 22. The two-stage thrust type rocket motor according to any one of the first to third inventions.

According to the first invention, since the second ignition casing loaded with the second igniting agent is disposed in the hollow portion of the second propellant, the second space for loading the second igniting agent is provided in the second space. There is no need to provide it in the front part of the combustion chamber. As a result, an increase in the structural weight due to an increase in the front structure of the combustion chamber can be prevented, the weight of the motor can be reduced, and the range can be extended.
In addition, the ignition gas of the second igniting agent is discharged from the through hole provided in the wall surface of the second ignition casing loaded and directly ignites the surface of the opposing second propellant, and at the same time, the ignition gas is The propellant surface that flows into the space between the diaphragm and the second propellant and faces the space can also be directly ignited. When this large area is directly ignited by the ignition gas of the second igniting agent, the initial generation amount of the combustion gas of the second propellant increases, and the time until the diaphragm breaks or disappears is shortened. Thereby, it is possible to prevent the second stage thrust from being delayed.
Thus, the flying object using the two-stage thrust type rocket motor of the present invention can obtain a reliable flying performance without causing problems such as a reduction in speed.
In addition, by attaching the first initiator that ignites the first igniting agent to the same blockade as the second initiator that ignites the second igniting agent, problems such as disconnection occur in the initiator that ignites the first igniting agent. Even when replacement is necessary, the initiator can be easily replaced without disassembling the motor. In addition, it is possible to easily incorporate a safety device (a device that prevents the ignition gas from the initiator from igniting the igniting agent inadvertently) attached to the igniter with a rocket motor into the sealed body to which the initiator is attached.

According to the second aspect of the present invention, the gas generated from the first propellant and the combustion gas from the first propellant do not break or disappear, and the gas from the combustion gas from the second propellant and the gas generated from the second propellant surely breaks. Alternatively, it disappears and the second propellant does not ignite before the second igniter ignites.
Further, according to the third invention, it is possible to prevent the diaphragm from being deformed due to its own weight during storage or vibration during transportation, and the thickness and arrangement required for the diaphragm can be maintained. .

Further, according to the fourth aspect of the present invention, when ignition cannot be reliably performed only with the ignition gas of the initiator, or ignition is delayed, ignition of the ignition agent is surely performed without delay with the ignition gas of the auxiliary agent. It becomes possible.

Embodiments of the present invention will be described below with reference to the drawings.
1 to 5 are diagrams showing an example of an embodiment of the reference invention, and FIGS. 6 to 8 are diagrams showing an example of the present invention.
First, FIG. 1 is an assembly view showing an example of a 2-stage thrust rocket motor of that. Figure 2 is an external view of a front assembly of the membrane used therefor, Fig. 3, the first combustion chamber to be used therefor is an enlarged view showing the connecting portion of the second combustion chamber and fittings, the 4 Figure is an enlarged view showing the igniter used therein, FIG. 5 is an enlarged view showing the another embodiment.
FIG. 6 is an assembly view showing another embodiment of the two-stage thrust rocket motor of the present invention, and FIGS. 7 and 8 are enlarged views showing an igniter used therefor.

1 to 5, the first combustion chamber 1 and the second combustion chamber 2 are connected in series by coupling of a joint 3 and a bolt. An injection nozzle 4 is attached to the rear end of the first combustion chamber 1. A heat insulating material 5 is attached to the inner surfaces of the first combustion chamber 1, the second combustion chamber 2 and the joint 3, and a first propellant 6 having a hollow portion penetrating in the front-rear direction and a hollow portion penetrating in the front-rear direction are provided. The 2nd propellant 7 which has is each arrange | positioned.
The igniter 8 in which the first ignition housing 12 and the second ignition housing 13 are connected in series is attached to the front portion of the second combustion chamber 2 with a bolt.

Further, in the case of this example, between the first propellant 6 and the second propellant 7, the outer peripheral end is the joint between the second combustion chamber 2 and the joint 3 on the entire circumference near the center of the rocket motor body. A diaphragm 9 fixed and having an inner peripheral end fixed between the first ignition housing 12 and the second ignition housing 13 separates the first propellant 6 and the second propellant 7 from each other. It is arranged.
The outer peripheral end of the diaphragm 9 may be fixed at the joint between the first combustion chamber 1 and the second combustion chamber 2 or at the joint between the first combustion chamber 1 and the joint 3.
Further, the diaphragm 9 is not broken or lost by the generated gas of the first igniter 10 and the combustion gas of the first propellant 6, but is broken by the combustion gas of the second propellant 7 and the generated gas of the second igniter 11. Alternatively, it must disappear so that the second propellant 7 does not ignite before the second igniter 11 ignites.

Such a diaphragm 9 has a shape as shown in FIG. 2, and its composition is such as EPDM rubber or silicone rubber containing organic fibers such as EPDM rubber, silicone rubber, Kevlar fiber, carbon fiber, etc. Silicone rubber or EPDM rubber containing inorganic fibers can be applied.
Among these, organic fiber or inorganic fiber-containing silicone rubber which is difficult to burn and disappear with the combustion gas of the first propellant and has high strength is preferable as the diaphragm composition. The shape as shown in FIG. 2 is formed by press molding, for example.

FIG. 3 shows a connecting portion of the first combustion chamber 1, the second combustion chamber 2 and the joint 3. The first combustion chamber 1 and the joint 3, and the second combustion chamber 2 and the joint 3 are respectively bolted. Are combined. When the second combustion chamber 2 and the joint 3 are coupled, the diaphragm 9 shown in FIG. 2 is fitted in the groove portion of the second combustion chamber 2 with its outer peripheral end fitted into the second combustion chamber 2 and the joint 3. It is attached in a state where it is pressed from both sides.
At this time, an adhesive can be applied and bonded to the contact portion in order to ensure the airtightness of the joint portion and the diaphragm 9 and the contact portion between the diaphragm 9 and the second combustion chamber 2. For example, when the diaphragm composition is a silicone rubber containing organic fibers or inorganic fibers, a silicone-based adhesive is applied.

FIG. 4 is an enlarged view of the igniter 8.
In the figure, the first igniting agent 10 is an igniting agent that selectively ignites the first propellant 6 and the second igniting agent 11 selectively ignites the second propellant 7. A first ignition case 12 and a second ignition case 13 each having a plurality of through-holes on the wall surface and made of a metal such as aluminum are loaded with igniting agents. The housing 12 is connected to the first sealing body 14 and the second ignition housing 13 is connected to the second sealing body 15 by screws.

  A first initiator 16 is attached to the first sealing body 14, and a second initiator 17 is attached to the second sealing body 15, and the second sealing body 15 is coupled to the front portion of the second combustion chamber 2 with bolts. The one sealing body 14 is coupled to the second ignition housing 13 with a screw.

  The leg of the first initiator 16 is mechanically and electrically coupled by, for example, a crimp terminal 50. The leg 18 connected to the crimp terminal 50 passes through the second ignition housing 13 and is second from the hole of the second sealing body 15 formed at the front end of the second ignition housing 13. It is taken out of the ignition housing 13. The hermeticity is maintained between the leg 18 and the second sealed body 15 by, for example, a glass hermetic seal 19.

Further, when the first sealing body 14 and the second ignition housing 13 are coupled by screws, the diaphragm 9 is sandwiched between the first sealing body 14 and the second ignition housing 13 and is pressed from both sides. The inner peripheral end is fixed in the state.
At this time, the contact portion can be adhered with an adhesive to ensure the airtightness of the contact portion between the diaphragm 9 and the first sealing body 14 and the contact portion between the diaphragm 9 and the second ignition housing 13. For example, when the diaphragm composition is a silicone rubber containing organic fibers or inorganic fibers, a silicone-based adhesive is applied.

Next, a method for setting the composition, thickness, and position of the diaphragm 9 will be described. The diaphragm 9 is deformed in the space formed by the second ignition housing 13 and the second propellant 7 due to the combustion gas pressure of the first propellant 6. The deformation of the diaphragm 9 is suppressed by the repulsive action against the deformation of the medicine 7. At this time, as the initial deformation of the diaphragm 9 increases, the tensile stress generated inside the diaphragm 9 increases, and when this tensile stress exceeds the allowable breaking stress of the material of the diaphragm 9, the diaphragm 9 breaks at that time. become.
Based on this principle, the outer shape of the second ignition housing 13 and the second propellant that determine the composition and thickness of the diaphragm 9 and the space in which the diaphragm 9 is deformed so that the tensile stress is less than the allowable breaking stress of the diaphragm 9. The arrangement of the diaphragm 9 with respect to the inner surface shape 7 is set.
The tensile stress generated in the diaphragm 9 necessary for setting can be obtained by calculation using a conventionally used stress analysis program.

In order to limit the size of the space, as shown in FIG. 5, the housing attached to the second ignition housing 13 by screw connection as a part of the second ignition housing 13 as necessary. The form which fixes the diaphragm 9 with the structure 20 of the same material as a body and the 1st blockade 14 may be sufficient.
By the above method, it is possible to use the diaphragm 9 that does not break even when subjected to the combustion gas pressure of the first propellant 6.

  The composition, thickness, and arrangement of the diaphragm 9 need to be set so that the temperature of the surface on the second propellant 7 side is lower than the ignition temperature of the second propellant 7 until a predetermined time when the second propellant 7 is ignited. There is. The composition, thickness, and arrangement of the diaphragm 9 are conventionally used so that the surface of the diaphragm 9 on the second propellant 7 side is below the ignition temperature until a predetermined time when the second propellant 7 is ignited. Can be set by program calculation.

  The composition, thickness, and arrangement of the diaphragm 9 in this example are set so as to satisfy both the conditions that do not break and the temperature conditions as described above. For example, as a first propellant 6 and a second propellant 7 having an outer diameter of 200 mm and a length of 2000 mm, a two-stage thrust rocket using a composite propellant mainly composed of a polybutadiene binder and ammonium perchlorate. In the case of a motor, the thickness of the diaphragm is set to a value between 1 mm and 20 mm.

As described above, the two-stage thrust type rocket motor of the first example is completed. However, depending on the thickness, composition, and arrangement of the diaphragm 9, the first combustion chamber is caused by its own weight during storage or vibration generated during transportation. There is a case in which the thickness and arrangement of the set diaphragm 9 are changed due to the large deformation to the 6 side.
In order to prevent this, the surface of the second propellant 7, that is, the rear end surface of the second propellant 7 or the second propulsion is used as necessary so that the diaphragm 9 does not greatly deform toward the first combustion chamber 1. Adhesion is also performed on a part or the entire surface of the diaphragm 9 on the surface of the hollow portion wall surface of the medicine 7 on the first combustion chamber side.
In this case, an adhesive is applied to the surface of the second propellant 7, and after the diaphragm 9 is adhered to the second propellant 7, the inner peripheral end of the diaphragm 9 is connected to the first sealing body 14 and the second casing for ignition. 13 to fix.

  The adhesive that bonds the diaphragm 9 and the second propellant 7 is selected according to the composition of the second propellant 7 and the diaphragm 9. For example, if the second propellant 7 is a composite propellant mainly composed of a polybutadiene binder and ammonium perchlorate and EPDM rubber containing Kevlar is used as the diaphragm 9, the adhesive may be, for example, an epoxy An adhesive is selected.

In addition, depending on the size of the two-stage thrust rocket motor, the amounts of the first and second igniters may increase, and the igniter may not be reliably ignited with only the ignition gas from the initiator. . In that case, it is possible to reliably ignite the igniting agent and eventually the propellant by arranging an auxiliary agent that ignites with the ignition gas of the initiator and ignites the igniting agent in the sealed body. It becomes.

Next, the operation of this example will be described.
When the first initiator 16 is energized, the first propellant 6 starts to burn by the ignition gas of the first igniter 10. At this time, the diaphragm 9 is deformed to the second combustion chamber 2 side by the generated combustion gas pressure of the first propellant 6, but the tensile stress generated in the diaphragm 9 is set to be equal to or less than the allowable breaking stress of the diaphragm material. Therefore, the first propellant 6 does not break during combustion. Further, since the temperature of the surface on the second propellant 7 side is set to be lower than the ignition temperature of the second propellant 7 by the diaphragm 9, the second propellant 7 is ignited during the combustion of the first propellant 6. Never do. By these actions, a first stage thrust is generated by the combustion of the first propellant 6.
After the combustion of the first propellant 6 is finished and the first stage thrust is finished, the surface of the diaphragm 9 on the second propellant 7 side of the diaphragm 9 until the predetermined time for generating the second stage thrust is Since the temperature is set to be lower than the ignition temperature, the second propellant 7 does not ignite and start combustion until that time. That is, the second stage thrust does not occur at an unexpected time.

When the second initiator 17 is energized at a predetermined time, the second igniter 11 is ignited, the ignited gas is released from the through-hole of the wall surface of the second ignition housing 13, and the second propellant 7 is Ignite.
Due to the combustion gas pressure generated by the second propellant 7, the diaphragm 9 is greatly deformed into the space of the first combustion chamber 1 where the first propellant 6 is eliminated. Due to this deformation, the diaphragm 9 is broken when the tensile stress generated inside becomes equal to or greater than the allowable breaking stress. Alternatively, the diaphragm 9 disappears by the heat of the combustion gas of the second igniter 11 and the second propellant 7 before breaking. The combustion gas of the second propellant 7 flows into the first combustion chamber from the broken or disappeared portion, and is then injected from the injection nozzle 4 to generate a second stage thrust.
At this time, since the ignition gas of the second propellant 11 ignites a wide area of the second propellant 7 at an initial stage, the generation rate of the combustion gas amount of the second propellant 7 is increased. That is, since the pressure rises faster, the time until the diaphragm 9 breaks is shortened, and the rise of the second stage thrust can be achieved in a short time after the initiator 17 is energized. Thereafter, the second stage thrust is generated until the combustion of the second propellant 7 is completed.

FIG. 6 is an assembly drawing of a two-stage thrust type rocket motor showing another example of the present invention. Since this example is different from the first example only in the structure of the igniter and the other structures are the same, the igniter will be described below.
FIG. 7 is an enlarged view around the igniter. In this example, the first igniting agent 10 and the second igniting agent 11 have one or more through holes on the wall surface. For example, the first igniting case 12 and the second igniting case 13 made of metal such as aluminum are used. Each is loaded. The second ignition housing 13 is coupled to a second sealing body 15 having a flange that can be coupled to the front portion of the second combustion chamber 2 with a bolt. The second sealing body 15 is loaded with an auxiliary medicine 21 for igniting the second ignition powder 11 in a circumferential groove, and an initiator 22 for igniting the auxiliary medicine 21 is attached.

FIG. 8 is a partially enlarged view of the igniter.
A tube 23 made of metal such as aluminum, for example, loaded with an auxiliary medicine 24 for igniting the first igniting agent 10 is connected to the second sealing body 15 with a screw, and the auxiliary medicine 24 is ignited. An initiator 26 is attached. The tubular body 23 penetrates through the second ignition casing 13 loaded with the second ignition powder 11, and the front portion thereof is located in the first ignition casing 12 loaded with the first ignition powder 10. It is attached as follows. The tube body 23 has a plurality of through holes in the wall surface located in the first ignition housing 12. In order to prevent unexpected ignition of the second igniter 11 due to the inflow of the ignition gas of the first igniter 10 and the combustion gas of the first propellant 6 into the second ignition housing 13 when the first stage thrust is generated, An o-ring 25 such as nitrile rubber, fluorine rubber, or silicon rubber is attached to the fitting portion between the first sealing body 14 and the tube body 23. First ignition housing 12 and first sealing body 14, second ignition housing 13 and second sealing body 15, first sealing body 14 and second ignition housing 13 coupling structure, and diaphragm 9 fixing method Is the same as in the first example.

Next, the operation of this example will be described.
When the initiator 26 is energized and ignites, the auxiliary medicine 24 ignites. The ignition gas of the assisting agent 24 is ejected from a through-hole in the wall surface of the tubular body 23 located in the first ignition housing 12 to ignite the first ignition agent 10, and the ignition gas is the first ignition housing 12. The first propellant 6 is ignited by igniting the first propellant 6 from the through-hole in the wall surface, and starts to burn. That is, the first stage thrust is generated. After the combustion of the first propellant 6 is completed, the initiator 22 is energized at a predetermined time, the assisting agent 21 is ignited by the ignition gas, and the second ignition agent 11 is ignited by the ignition gas. The ignition gas is ejected from a through-hole in the wall surface of the second ignition housing 13 to ignite the second propellant 7, and the diaphragm 9 is broken or disappeared by the pressure or heat of the combustion gas. The combustion gas of the propellant 7 flows into the first combustion chamber 1 and is then injected from the injection nozzle 4 to generate a second stage thrust.

As explained above, according to the present invention, in the state where the first propellant is burning, the diaphragm disposed between the second propellant and the first propellant does not break and disappear, Furthermore, since the surface temperature of the second propellant side of the diaphragm is kept below the ignition temperature of the second propellant until a predetermined time at which the second propellant is ignited, the second propellant starts to combust until then. There is nothing.
Next, the second igniter is ignited, and the second propellant is ignited by the ignited gas and combustion is started. At this time, the diaphragm breaks or disappears due to the pressure of the ignition gas of the second igniting agent and the combustion gas of the second propellant, or heat, and the combustion gas of the second propellant flows into the first combustion chamber from that portion, Sprayed from the spray nozzle. Thus, a two-stage thrust type rocket motor is obtained in which the first stage thrust is generated by the combustion of the first propellant and the second stage thrust is generated by the combustion of the second propellant that is started at a predetermined time.

It is an assembly drawing which shows an example of the two-stage thrust type rocket motor of the reference invention. It is an external view which shows the pre-assembly of the diaphragm used in FIG. It is an enlarged view which shows the connection part of the 1st combustion chamber of FIG. 1, a 2nd combustion chamber, and a joint. It is an enlarged view which shows the igniter of FIG. It is an enlarged view which shows another form of the igniter of FIG. It is an assembly drawing which shows other embodiment of the two-stage thrust type rocket motor of this invention. It is an enlarged view which shows the igniter periphery of FIG. It is an enlarged view which shows a part of igniter of FIG. It is an assembly drawing of a two-stage thrust type rocket motor showing the prior art. It is an enlarged view which shows the igniter attachment part of FIG.

Explanation of symbols

  1: first combustion chamber, 2: second combustion chamber, 3: joint, 4: injection nozzle, 5: heat insulating material, 6: first propellant, 7: second propellant, 8: igniter, 9: diaphragm, 10: first igniter, 11: second igniter, 12: first ignition housing, 13: second ignition housing, 14: first sealing body, 15: second sealing body, 16: first Initiator, 17: second initiator, 18: leg wire, 19: glass hermetic seal, 20: structure, 21: second assisting agent, 22: second initiator, 23: tubular body, 24: first assisting agent , 25: O-ring, 26: first initiator, 30: injection nozzle, 31: Porlabus, 32: Porlabus, 33: combustion chamber, 34: first propellant, 35: second propellant, 36: forward insulator, 37: diaphragm, 38: propellant support, 39: igniter blockade, 40: igniter case 41: first igniter, 42: second igniter, 43: annular chamber, 50: crimp terminal, 51: a nozzle port.

Claims (4)

An injection nozzle 4 is arranged at the rear,
The first combustion chamber 1 and the second combustion chamber 2 are disposed in front of the injection nozzle 4 in such a state that the first combustion chamber 1 and the second combustion chamber 2 are connected in series in this order directly or via the joint 3.
A first propellant 6 having a hollow portion penetrating front and rear in the first combustion chamber and a second propellant 7 having a hollow portion penetrating front and rear in the second combustion chamber, respectively;
One or more through holes are provided in the wall surface of the position where the first propellant 6 is selectively ignited, and a mechanism for igniting the first igniter 10 is attached to the front end of the first ignition housing 12. A first ignition housing 12 having a first sealing body 14 and loaded with a first igniting agent 10 is disposed on a front and rear central axis in the first combustion chamber,
One or more through holes are provided in the wall surface at a position where the hollow wall surface of the second propellant 7 is directly ignited, and a mechanism for igniting the second igniter 11 is attached to the front end of the second ignition housing 13. A second ignition housing 13 having the second sealing body 15 and the second ignition powder 11 loaded therein is disposed in the hollow portion of the second propellant 7;
A diaphragm 9 having an outer peripheral end fixed on the entire circumference of the rocket motor body and an inner peripheral end fixed between the first ignition housing 12 and the second ignition housing 13 is a first propellant. 6 and the second propellant 7 are disposed to be isolated from each other,
The diaphragm 9 is not broken or lost by the gas generated from the first igniter 10 and the combustion gas from the first propellant 6, but is broken or lost by the combustion gas from the second propellant 7 and the gas generated from the second igniter 11. In the two-stage thrust type rocket motor, the mechanism for igniting the first igniting agent 10 includes all of the first initiator 26, the second sealing body 15, the second ignition housing 13 and the first sealing body 14 connected thereto. A two-stage thrust rocket characterized by comprising a tubular body 23 that penetrates and extends to the inside of the first ignition housing 12 and that has an auxiliary agent 24 that propagates ignition of the initiator 26 inside. motor. The two-stage thrust type rocket motor according to claim 1, wherein the diaphragm 9 is EPDM rubber, silicone rubber, EPDM rubber or silicone rubber containing organic fibers, or EPDM rubber or silicone rubber containing inorganic fibers. The two-stage thrust rocket motor according to any one of claims 1 to 2, wherein the diaphragm (9) is bonded to the rear end face of the second propellant (7) or the hollow portion wall surface of the second propellant (7). . The mechanism for igniting the second igniting agent 11 includes an initiator 22 and a second auxiliary agent 21 in the second sealing body 15 that propagates the ignition of the initiator 22. The two-stage thrust type rocket motor according to any one of the above.

Images