Description:FIELD OF THE INVENTION:

The present invention generally relates to the field of aerospace. Particularly, the invention relates to the propulsion system for airborne such as missile, rocket or the like. More particularly, the invention relates to the unified propulsion system for the Anti-Tank Guided Missile (ATGM).

BACKGROUND OF THE INVENTION:

Propulsion system is an important sub-system in any missile. The propulsion system is used to produce thrust which is used to moves the rocket/missile through the forward direction. The propulsion system consists of solid propellant and especially required for initiating the launch, to boost or sustain the missiles. Generally, the rocket motor comprises a metallic cylindrical casing along with insulation (safeguard the metallic casing from hot gases), called as combustion chamber. The said combustion chamber is filled with propellant which upon ignition and burning generates large volumes of gases at high pressures and temperatures.

Propulsion system provides the required thrust for the missile to eject from the launch tube and to travel towards the target. There are many types of propulsion system described in the prior arts. Some existing prior arts are as follows:

US patent no. US9169806B2 discloses propulsion system for flying machine, particularly for a missile. The propulsion system includes a booster, a turbojet engine, the booster including a chamber which is fixed to a rear casing of the turbojet engine by being arranged along the longitudinal axis thereof. The chamber includes at a rear a jet pipe and includes at least one charge and a mechanism initiating the charge. The gas bleed tubes connected to the booster, and which are configured either for igniting the combustion chamber of the turbojet engine on for starting the turbine of the turbojet.

German patent no. DE102016121081B4 discloses ejection engine as an annular combustion chamber with a device for stabilizing a propellant arrangement with the formation of outflow channels. It includes a missile which comprises an aircraft engine with at least one combustion chamber of the aircraft engine and with at least one exhaust engine for generating a starting thrust of the aircraft engine in a starting phase. The ejection engine an annular combustor having disposed therein at least one annular propellant having at least one molded structure integral with at least one propellant assembly, said annular propellant having at least one effluent duct and the at least one mold structure and the at least one propellant assembly are spiral or volute shaped and the mold structure has a profile and the profile is wavy and the wavy mold structure inboard to the inner lateral surface or on the outside to the outer lateral surface of the strip-shaped or plate-shaped propellant arrangement.

The drawback in the above existing prior arts is that the leakage of the gases produced in the launch side. The developed pressure and higher temperature generated during the combustion of propellant are affected the casing of the missile/rocket. Although the above prior arts provide propulsion system for missile there still needs exist to develop a unified propulsion system to overcome the above drawbacks mentioned in the existing prior arts. The system should withstand to the pressure developed during the combustion of the propellant throughout the flight time. It should be leak proof and should withstand higher temperatures generated during the combustion of the propellant. Insulation should be provided to minimize the temperature on the outside surfaces to protect other integrated systems. The system should have provision for Igniter to initiate the ignition process and Pressure Transducer for measuring the pressure developed inside the chamber.

Thus, the present invention discloses a unified propulsion system for the aforementioned purpose which provides a leak proof propulsion system. The propulsion system withstands to the developed pressure and generated temperature during the combustion of the propellant.

Thus, the aim of present invention is to solve all these issues of the known prior arts, by aiming to provide a unified propulsion system for Anti-Tank Guided Missile (ATGM).

OBJECTIVES OF THE INVENTION:

The principal object of the present invention is to provide a unified propulsion system for Anti-Tank Guided Missile (ATGM).

Another main object of the present invention is to provide a unified propulsion system which reduces the requirement of multiple components and uses a single nozzle for both the purposes i.e. for launch as well as flight motor.

Another object of the invention is to provide a unified propulsion system having a leakproof property and also to withstand higher temperatures generated during the combustion of the propellant.

Another object of the invention is to provide insulation in the propulsion system to minimize the temperature on the outside surfaces to protect other integrated systems.

Another object of the invention is to provide an igniter to ignite the ignition process and a pressure transducer to measure the pressure developed inside the chamber of the propulsion system.

Another object of the invention is to provide a simplified design of unified propulsion system which is used for different applications as per the requirement with minor modifications.

Yet another object of the invention is to provide the whole propulsion system as compact designed single unit and also make it easy to handle and use.

SUMMARY OF THE INVENTION:

Accordingly, the present invention provides a unified propulsion system for Anti-Tank Guided Missile (ATGM). The unified propulsion system typically consists of flight motor assembly and launch motor assembly along with nozzle assembly and other components. These two sub-assemblies i.e., flight motor assembly and launch motor assembly are joined together with fasteners. Further, these sub-assemblies are provided with sealing to protect the leakage of hot gases produced in the launch side, and thereby loss of developed thrust.

The propulsion system is designed to withstand to the pressure developed during the combustion of the propellant throughout the flight time. It provides a leak proof property and withstand higher temperatures generated during the combustion of the propellant. Insulation minimizes the temperature on the outside surfaces to protect other integrated systems. The propulsion system comprises an igniter to initiate the ignition process and a pressure transducer for measuring the pressure developed inside the chamber.

In one aspect of the present invention, the present invention provides a unified propulsion system (100) for Anti-Tank Guided Missile (ATGM). The said unified propulsion system (100) comprising:
- a flight motor assembly (A);
- a launch motor assembly (B); and
- a nozzle assembly (C),
wherein the unified propulsion system (100) is provided with single nozzle assembly (C) and an integrated combustion chamber for both purpose launch and flight motor;
wherein the rear end of the flight motor assembly (A) is connected with the launch motor assembly (B) and the nozzle assembly (C); and
wherein the said the flight motor assembly (A) and the launch motor assembly (B) are joined together with fasteners.

The flight motor assembly (A) comprising:
- flight motor grain (1);
- a front cover (2)
- sealings (3)
- a flight motor body (4);
- an outer shell (5); and
- a pressure transducer (6);
wherein the flight motor grain (1) is a solid propellent and a metallic housing is provided in flight motor assembly (A) for the flight motor grain (1);
wherein the front cover (2) is designed to close the combustion chamber; and
wherein the sealings (3) are provided to protect the leakage of hot gases produced in the combustion chamber.

The launch motor assembly (B) comprises:
- an igniter (7);
- sealing (8);
- a grid support (9);
- a washer (10);
- a grid separator (11);
- a blast tube (12); and
- hollow propellant sticks (13);
wherein the igniter (7) is provided in the blast tube (12) to initiate the combustion of hollow propellant sticks (13); and
wherein the said grid support (9) is provided to support the grid separator (11) to prevent the damage during the launch motor combustion.

The said nozzle assembly (C) comprising:
- an INSITU molded rear grid (14); and
- a nozzle (15);
wherein the nozzle (15) exhausts the produced gases from the launch side; and
wherein the INSITU molded rear grid (14) protects the nozzle (15) from blockage of unburnt propellant chips and withstand high temperature and high pressure of gases.

The said pressure transducer (6) monitors the pressure produced during combustion process.

The said igniter (7) is fixed in the projection provided at an angle to the axis of the fight motor assembly (A).

The sealings (3, 8) are provided in the flight motor assembly (A) and launch motor assembly (B) to protect the leakage of hot gases produced in the combustion chamber.

The washer (10) is provided to protect leakage of hot gases and withstand high temperatures;
wherein during operation of leak check, the washer (10) separates the flight motor assembly (A) and launch motor assembly (B).

Thus, the aim of the present invention is to develop and design a novel unified propulsion system (100) for Anti-Tank Guided Missile (ATGM).

The above description merely is an outline of the technical solution of the present disclosure; in order to know the technical means of the present disclosure more clearly so that implementation may be carried out according to contents of the specification, and in order to make the above and other objectives, characteristics and advantages of the present disclosure more clear and easy to understand, specific embodiments of the present invention will be described in detail below.

BRIEF DESCRIPTION OF DRAWINGS:

The accompanying drawing which is incorporated herein constitute a portion of this specification and illustrate exemplary practices according to the invention which, together with the general description above and the detailed description set forth below will serve to explain the principle of the invention wherein:

Fig. 1 shows the schematic perspective view of unified propulsion system (100) for Anti-Tank Guided Missile (ATGM) of the present invention.

DETAILED DESCRIPTION OF THE INVENTION:

Accordingly, the present invention provides a unified propulsion system for Anti-Tank Guided Missile. The propulsion system of rocket or missile is used to produce thrust. The produced thrust is force which moves the rocket or missile in a forward direction. In general, the propulsion system includes combustion chamber, nozzle and igniter. The reaction of propellent takes place in the chamber and produce gases. The produced gases have very high temperature and pressure.

The propulsion system provides the required thrust for the missile to eject from the launch tube and to travel towards the target. The system should withstand to the pressure developed during the combustion of the propellant throughout the flight time. It should be leak proof and should withstand higher temperatures generated during the combustion of the propellant. Insulation should be provided to minimize the temperature on the outside surfaces to protect other integrated systems.

Therefore, present invention provides a unified propulsion system for the aforementioned purpose. The said unified propulsion system typically consists of flight motor assembly and launch motor assembly along with nozzle assembly and other components. These two sub-Assemblies i.e., flight motor assembly and launch motor assembly are joined together with fasteners and sealing is provided to protect the leakage of hot gases and thereby loss of developed thrust.

The said propulsion system reduces the requirement of multiple components and make use of single nozzle for both the purposes (launch as well as flight motor). It is a simple design and can be used in different applications as per the requirement with minor modifications. The system includes an igniter to initiate the ignition process and a pressure transducer for measuring the pressure developed inside the combustion chamber.

In one aspect, the present invention provides a unified propulsion system for Anti-Tank Guided Missile (ATGM). The propulsion system provides a compact single unit design which is easy to handle and use. The said propulsion system divided into three major parts such as Flight motor assembly, Launch motor assembly and Nozzle assembly.

Fig. 1 shows the Anti-Tank Guided Missile (ATGM) with the unified propulsion system (100). The said ATGM comprises a Flight motor assembly (A), a Launch motor assembly (B), a Nozzle assembly (C). The said assemblies are further comprises a flight motor grain (1), a front cover (2), sealings (3), a motor body (4), an outer shell (5), a pressure transducer (6), a flight motor igniter (7), sealing (8), a grid support (9), a washer (10), a grid separator (11), a blast tube (12), propulsion sticks (13), an INSITU molded rear grid (14) and a nozzle (15).

The propulsion system (100) of ATGM comprises three main assemblies i.e. a Flight motor assembly (A), a Launch motor assembly (B) and a Nozzle assembly (C). The said propulsion system (100) uses single nozzle (15) for both the purpose launch as well as flight motor. The said two sub-assemblies flight motor assembly (A) and launch motor assembly (B) are joined together with fasteners. Further, sealings are provided in sub-assemblies i.e., flight motor assembly (A) and launch motor assembly (B) to protect the leakage of hot gases produced during the combustion process.

The novel propulsion system is provided with an integrated chamber for both flight motor as well as launch motor. Combustion for both launch and flight motors take place in the said integrated chamber.

Flight Motor Assembly (A) comprises a motor body (4) and an outer shell (5) for the motor body (4). A housing is provided in flight motor assembly (A) for solid propellant for the combustion process. In the housing the said solid propellants are in the form of solid grains (1) which are burnt once the ignition process starts. In one embodiment, the flight motor assembly (A) is designed to provide metallic housing for solid propellant.

Flight motor assembly (A) comprises a front cover (2) which is designed to close the chamber. In one embodiment, the said front cover (2) is in dome shaped. Sealings (3) are provided at surrounding the housing or chamber of flight motor assembly (A) to provide a leak proof technology and withstand the pressure developed within the combustion chamber. The said sealings (3) prevents leakage of hot pressurized gases and unburnt propellant from the flight motor assembly (A).

The design of the flight motor assembly (A) is completely sealed to withstand a pressure generated by burnt gases inside the combustion chamber.

The igniter (7) is assembled in the flight motor assembly to achieve the task. The igniter (7) ignites the solid propellant present in the combustion chamber. The surface of solid propellent grain burns and produces hot gas, which is expelled from the combustion chamber. The propulsive force of a solid propellant motor is derived from the combustion of solid propellant at high temperature and pressure. The igniter (7) induces the combustion reaction in a controlled and predictable manner by generating heat flux in the form of hot dense gases that rapidly ignite the propellant surface.

The igniter (7) also contributes towards the generation of a certain minimum pressure inside the motor that is adequate for stable and sustained combustion of the propellant. Igniter (7) is provided with sealing (8) which ensures sealing of gases. In one embodiment, the igniter (7) is fixed in the projection provided at an angle to the axis of the assembly (A). Further, insulation is provided to protect the metallic casing from high temperature gases during combustion of the propellant.

The pressure transducer (6) is assembled in the flight motor assembly (A). The pressure transducer (6) consists of a pressure-sensitive element, such as a diaphragm, with a constant area. The fluid pressure causes the diaphragm to deflect. The pressure transducer (6) also consists of a transduction element. This transduction element converts the deflection sensed by the diaphragm into an electrical output signal. This signal increases or decreases proportionally to the pressure change. Therefore, device calibration is critical to ensure that the pressure is within the range of the specifications. Provision for placing the pressure transducer (6) is provided to monitor the pressure produced during combustion in the combustion chamber.

Further, the rear side of the flight motor assembly (A) is to connect the extensions like Launch Motor Assembly (B) and Nozzle Assembly (C) to complete the combustion process.

Launch Motor Assembly (B) is designed to provide a housing for hollow propellant sticks (13). The said housing is used to withstand a very high pressure. In one embodiment, the launch motor assembly (B) having a metallic housing for hollow propellant sticks (13). Hollow propellant sticks (13) provide very high burn rate because of high surface area, hence it provides high thrust to launch the missile within few milliseconds.

The launch motor assembly (B) is provided with insulation to protect the metallic casing of launch motor assembly (B) from high temperature of gases during combustion of the propellant. The integrated chamber is used in the Anti-Tank Guided Missile for both the flight motor and launch motor. The launch motor insulation is a critical design as the gases during launch motor combustion and flight Motor combustion passes through the same chamber.

Launch motor assembly (B) consists of a blast tube (12), grid support (9), grid separator (11) and propellant sticks assembly (13).

Igniter (7) is provided in the blast tube (12) to initiate the combustion of propellant sticks (13). Grid separator (11) is provided to protect self-ignition of flight motor grain (1) from the hot gases generated during the combustion of launch motor propellant.

Grid Support (9) is provided to support the grid separator (11) to prevent the damage during the launch motor combustion. Dowel pins are provided to stop the rotation of grid separator (11). Annular holes are provided to allow the hot gases from flight motor while discharging the gases with minimum losses. The component is in direct contact with the hot gases which needs to withstand from melting. The material is also chosen after going through many aspects and optimum working pressure.

Washer (10) is provided to protect leakage of hot gases and withstand high temperatures. With the help of this washer (10), the launch motor is separated from flight motor during leak checks and thus launch motor assembly procedure is divided from Flight Motor.

The propulsion system uses single Nozzle Assembly (C) for both flight and launch motor. It is a metal-nonmetal molded assembly to get maximum output with minimum weight. Nozzle assembly (C) design is a critical aspect as the required thrust to be developed during the combustion of both launch motor and flight motor. The pressure to be converted to velocity with innovative design of convergent, throat and divergent portions with space constraint.

INSITU molded rear grid (14) is also critical component to include a metallic part in between non-metallic component. Rear grid (14) is provided to protect nozzle (15) from blockage of unburnt propellant chips, as well as withstand high temperature and high-pressure gases also and proper passage to flow the gases.

ADVANTAGES AND APPLICATION OF THE INVENTION:

This unified propulsion system has following advantageous features:
a) It has an integrated combustion chamber for both launch motor and flight motor.
b) The propulsion system uses single nozzle for complete operation.
c) It has optimized design for required thrust.
d) Easy to manufacture and assemble, no internal milling operation and welding operations are required.
e) Annular space around launch motor can be used for integration of other systems.
f) The propulsion system provides a leak proof technology for the combustion process. , Claims:
1. A unified propulsion system (100) for Anti-Tank Guided Missile, wherein the unified propulsion system (100) comprising:
- a flight motor assembly (A);
- a launch motor assembly (B); and
- a nozzle assembly (C),
wherein the unified propulsion system (100) is provided with single nozzle assembly (C) and an integrated combustion chamber for both purpose launch and flight motor;
wherein the rear end of the flight motor assembly (A) is connected with the launch motor assembly (B) and the nozzle assembly (C); and
wherein the said flight motor assembly (A) and the launch motor assembly (B) are joined together with fasteners.

2. The unified propulsion system (100) as claimed in claim 1, wherein the flight motor assembly (A) comprising:
- flight motor grain (1);
- a front cover (2);
- sealings (3);
- a flight motor body (4);
- an outer shell (5); and
- a pressure transducer (6);
wherein the flight motor grain (1) is a solid propellent and a metallic housing is provided in flight motor assembly (A) for the flight motor grain (1);
wherein the front cover (2) is designed to close the combustion chamber; and
wherein the sealings (3) are provided to protect the leakage of hot gases produced in the combustion chamber.

3. The unified propulsion system (100) as claimed in claim 1, wherein the launch motor assembly (B) comprises:
- an igniter (7);
- sealing (8);
- a grid support (9);
- a washer (10);
- a grid separator (11);
- a blast tube (12); and
- hollow propellant sticks (13);
wherein the igniter (7) is provided in the blast tube (12) to initiate the combustion of hollow propellant sticks (13); and
wherein the said grid support (9) is provided to support the grid separator (11) to prevent the damage during the launch motor combustion.

4. The unified propulsion system (100) as claimed in claim 1, wherein the said nozzle assembly (C) comprising:
- an INSITU molded rear grid (14); and
- a nozzle (15);
wherein the nozzle (15) exhausts the produced gases from the launch side; and
wherein the INSITU molded rear grid (14) protects the nozzle (15) from blockage of unburnt propellant chips and withstand high temperature and high pressure of gases.

5. The unified propulsion system (100) as claimed in claim 2, wherein the said pressure transducer (6) monitors the pressure produced during combustion process.

6. The unified propulsion system (100) as claimed in claim 3, wherein the said igniter (7) is fixed in the projection provided at an angle to the axis of the fight motor assembly (A).
7. The unified propulsion system (100) as claimed in claim 1, wherein sealings (3, 8) are provided in the flight motor assembly (A) and launch motor assembly (B) to protect the leakage of hot gases produced in the combustion chamber.

8. The unified propulsion system (100) as claimed in claim 3, wherein the washer (10) is provided to protect leakage of hot gases and withstand high temperatures;
wherein during operation of leak check, the washer (10) separates the flight motor assembly (A) and launch motor assembly (B).