RU2719799C1 - Signal feeder for separation of starting stage of vertically launching aircraft - Google Patents

Abstract

FIELD: missiles.

SUBSTANCE: invention relates to signal feeders for separation of starting stage of vertically launching aircraft with liquid-propellant engine and can be used in design and production of new rocket equipment. Proposed device consists of float. It is located with possibility of free movement along vertical axis inside the turned perforated guide sleeve. Said cup is fixed inside fuel tank at its bottom. There is a microswitch. Its housing is located outside the fuel tank. Rod of the microswitch is arranged with the possibility of free movement along the vertical axis in the lower part of the guide cup and is tightly brought out beyond the cavity of the fuel tank. Float in its lower position is provided with the possibility of pressing the microswitch rod, closing its electric contacts, signaling actual fuel generation from the tank and switching on the mechanisms of the start stage separation. Effort of float pressing on rod of microswitch is provided taking into account of flight overload.

EFFECT: technical result is higher reliability of device due to possibility of determining moment of separation of starting stage of vertically launching aircraft.

1 cl, 2 dwg

Description

The claimed invention relates to devices for supplying a signal to the separation of the starting stage of a vertically launching aircraft (LA) with a liquid rocket engine (LRE) and can be used in the design and manufacture of new rocket technology.

To determine the moment of separation of the LA launch stage, various devices are used. Basically, program-time devices are installed that take into account fuel consumption per unit time and give a command to divide the steps in accordance with the fuel consumption calculated in advance.

Known rocket with detachable missile part [1. Pat. 2406968 RU, IPC ⁶ F42B 15/00, F42B 15/36. Missile with detachable missile part / Makarovets N.A. et al. - Decl. 10/22/2009; publ. 12/20/2010, Bull. No. 35.], Containing a detachable missile part, a warhead and a powder separation charge disposed between them. At the front end of the rocket part before the separation charge, a damping unit is placed in the form of a set of six plate compensators of the same shape. After launch, the missile moves along a trajectory as close as possible to the calculated one, and at a given point in time, a separation charge is triggered at a given point in the trajectory.

Known multiple missile [2. Pat. 2522537 RU, IPC ⁶ F42B 12/58, F42B 15/36. Multiple Missile / Makarovets N.A. et al. - Decl. 03/13/2013; publ. 07/20/2014, Bull. No. 20.], Containing a rocket engine with a bottom, a detachable warhead, a parachute compartment, as well as an explosive device, in which an electronic temporary device with a safety-actuating mechanism, a reaction target sensor with a shock mechanism, and an actuating unit equipped with powder charges are placed. After the launch of the fissile shell, at the end of the counting of the pre-set time of the remote action, the electronic-temporary device of the fuse generates an electrical impulse and ensures the operation of the expelling charge.

The disadvantages of the above analogues [1., 2.]:

- software-temporary devices complicate the design of the aircraft, because to bring the systems into operation, an electrical connection with the control system and the presence of command lines in the program-temporary device for determining the necessary moment of operation are required;

- a large number of components difficult to manufacture increases the cost of the product.

In some cases, devices are used that include flowmeters that take into account the real fuel consumption and issue a separation command when the actual amount of fuel is drawn closer to the calculated minimum allowable volume. A known system of combined control of fuel consumption for a rocket propulsion system [3. Pat. 2685161 RU, IPC ⁶ F02K 9/56. Combined fuel management system for rocket propulsion / Kurdish S.S. - Declared. 07/18/2018; publ. 04/16/2019, Bull. No. 11.], Which contains sensors of the second fuel consumption: oxidizer and fuel, connected through an amplification-converting device to an on-board central computer (BCM), which supplies a control signal to the throttle actuator control unit, and fiber-optic level gauges connected to the BCM. In this case, the fiber-optic level gauges contain a cylindrical pipe with a distributed sensor located in it, communicating with a controlled fuel tank. In this case, the distributed sensor consists of an optical fiber, an optical reflectometer and an acting element - a float of magnetic material, inside of which three rollers are installed.

The disadvantages of this system:

- the presence of level gauges and computing devices interconnected electrically, leads to a complication of the design of the aircraft;

- a lot of complex devices and sensors increase the cost of the product;

- determination of fuel consumption occurs with an error, which leads to the need to have a guaranteed fuel supply, which will remain undeveloped.

Known rocket and the mechanism of separation of the stages of the rocket [4. Pat. 2284460 RU, IPC ⁶ F42B 15/00. Rocket and drive of the separation mechanism of rocket stages / Shipunov A.G. et al. - Decl. 02/08/2005; publ. 09/27/2006, Bull. No. 27.]. The missile contains a marching stage, a detachable starting stage with an engine, a stage separation mechanism including a drive, a prechamber with a pyro-composition and an igniter capsule, a transition frame. The rocket device implements a stage separation mechanism drive, which contains a piston and stoppers, when starting the engine of the starting rocket, gas from the combustion chamber enters the drive cavity through the holes and moves the piston, while the stoppers disengage from the housing and the gas pressure moves inside the piston. At the end of the starting engine, the pressure in the combustion chamber begins to drop, and the piston moves under the action of the installation spring until the ignition mechanism is triggered. The igniter capsule initiates pyrocomposition. The combustion products of the pyro-composition, expanding, move the prechamber to the stop in the engine, while the launch stage is disengaged.

The disadvantages of the analogue, reducing the reliability of the aircraft as a whole and the separation of stages in particular, are as follows:

- the design contains such mechanical elements that are difficult to position, such as a piston interacting with each other, stoppers, springs;

- premature operation of the system during fluctuations or pulsations of pressure in the combustion chamber is possible, since the design uses the gas pressure of the engine as the driving force of the drive and as the only factor and criterion for evaluating the optimal moment of separation of the stages.

Known devices that use to determine the time point of separation of a linear acceleration sensor. These include a missile [5. Pat. 2167388 RU, IPC F42B 15/10 (2000.01). A missile with a detachable engine / Dudka V.D. et al. - Decl. 02.24.2000; publ. 05/20/2001, Bull. No. 14.], Containing the marching stage and the engine docked to it with a stage separation mechanism. Parallel to the longitudinal axis of the projectile installed actuator remote inertial action. The mechanism is made in the form of a non-conductive housing with a fixed contact group and a conductive inertial body, tightened by a power spring in the direction of the stage separation mechanism. The inertial body is made in the form of a cylinder with an internal conical hole passing into the cylinder. The mass of the inertial body and the force of the power spring are calculated according to the dependence given in the claims, and provide a signal to the stage separation mechanism at the moment the projectile engine is finished, that is, the moment of fuel production is judged by the change in overload due to a drop in engine thrust.

A similar principle is applied in the separation device march stage LA [6. Pat. 2584401 RU, IPC ⁶ F42B 15/00. The method of separation of the march stage of the aircraft and a device for its implementation / Obrazumov V.I. et al. - Decl. 02/17/2015; publ. 05/20/2016, Bull. No. 14.]. The method includes mechanical holding in open state of the start circuit of the electric igniter of the separation mechanism of the steps when starting the aircraft at the start of the flight path. On board the aircraft during the flight, the longitudinal acceleration is measured and the rate of its change in time is determined, the moment of the sign change of the derivative of the longitudinal acceleration from negative to positive is recorded, and the electric igniter starting circuit of the stage separation mechanism is closed.

The aircraft contains a marching stage and an engine docked to it with a stage separation mechanism equipped with an electric igniter. In the march stage of the aircraft parallel to the longitudinal axis of the projectile installed actuator remote inertial action. The mechanism is made in the form of a current-conducting case with a fixed contact group and a conductive inertial body, locked by a power spring in the direction of the stage separation mechanism. The device also contains a linear acceleration sensor, an analog-to-digital converter and a computer.

The disadvantages of the analogues [5., 6.], reducing the reliability of the aircraft, the following:

- the complexity of structural elements such as a linear acceleration sensor, an analog-to-digital converter and a computer;

- the moment of fuel production is judged by the change in overload due to a drop in engine thrust, which can lead to a loss in aircraft speed.

The most significant drawback of the principles of operation of the signal supply devices for separation of the launch stage described in the analogues [1.-4., 6.] is that at the moment of separation of the launch stage of the aircraft there must be undeveloped fuel, which forces the aircraft to be filled with fuel in volume, exceeding the aircraft necessary for a normal flight, which is economically disadvantageous, as well as explosive at the time of the launch stage fall. At the same time, a flight with fully exhausted fuel is unacceptable, since with a later separation of the starting stage, the drag force will exceed the thrust force of the starting engine. An idle engine will slow down the aircraft, which will lose speed. Therefore, the command for separation should be given with a certain amount of fuel, i.e. with obviously untreated.

It is obvious that it is rational to separate the launch stage of an aircraft with an LRE not according to the instructions of a temporary device, flow meter or level gauge, nozzle pressure monitoring system and acceleration sensor, but according to the actual optimal fuel generation from the fuel tank. It is possible to control the production of liquid fuel without resorting to the complex devices mentioned above by using a float device.

There are various designs of floats and level sensors based on them for signaling.

A device for automatically indicating a critical oil level [7. Pat. 2874243 US, IPC G01F 23/30, G01F 23/64. Device for automatically indicating the critical oil level in engines, especially engines of motor vehicles / Metzler O. (DE) - Claim. 04/19/1957; publ. 02.17.1959.], In which a metal float moves freely in a metal pipe and closes the electric circuit when the oil level drops to a critical point.

An alarm sensor is also known [8. Pat. 1642434 US, IPC B60K 15/03, B60K 15/06, G01F 23/30, G01F 23/64. Alarm gauge / Raye FB - Declared. 10/02/1923; publ. 09/13/1927.], Containing a conductive ball-shaped float, closing the contacts of the housing in the lower position and thereby signaling the emptying of the tank.

The disadvantages of the devices [7., 8.] are as follows:

- conductive float closes electrical contact in the environment of fuel vapor, which can lead to sparking and fire or explosion;

- aggressive fuel can destroy conductive bare contacts.

Known liquid switch [9. US Pat 8,319,124, IPC H01H 35/18. Liquid activated switch apparatus / Wang N.S. - Declared. 06/28/2010; publ. 11.27.2012.], Which includes a container adapted to receive a liquid, such as water or rain. There are two electrically conductive elements inside the container, the first element attached to the container and the second to the floating body. The floating body is designed and configured so that it begins to rise along with an increase in the liquid level, which leads to the displacement of the second electrically conductive element that is attached to it, as a result of which two electrically conductive elements either close the electric circuit or open it.

The disadvantages of the device [9.] the following:

- the device is operable only at a positive ambient temperature;

- the controlled environment, water, is in the same cavity with exposed contacts, which is why a short circuit is possible during operation.

Known electric samovar [10. A.S. 1743576 SU, IPC A47G 19/14, NOSH 35/18 (1992.06). Electric Samovar / Sorokin V.P. - Declared. 01/16/1990; publ. 06/30/1992.], Having a housing, a heating element and voltage input devices located in the upper part of the housing and electrically connected to the closing contacts with the control button inserted into the float switch. The float is installed in the lower part of the housing with the possibility of influencing the control button through the introduced harmonic membrane, hermetically mounted in the upper part of the housing.

The disadvantage of the analogue [10.] is that the voltage input connectors are located in the upper part of the housing, which significantly increases the dimensions of the device.

The purpose of the invention is to improve the reliability of the system for determining the moment of separation of the launch stage of an aircraft with a rocket engine and to ensure automatism of its operation while reducing the number of intermediate control and actuating mechanisms.

The inventive signal supply device for separating the launch stage of a vertically launching aircraft with an LRE consists of a float, an inverted perforated guide cup and a microswitch with a rod. The guide cup is fixedly mounted inside the fuel tank, in its lower part. The float is freely movable along the vertical axis inside the guide cup. The microswitch housing is located outside the fuel tank. The microswitch rod is fixed with the possibility of movement along the vertical axis in the lower part of the guide glass and hermetically removed for the cavity of the fuel tank.

The inventive device is illustrated by drawings. In FIG. 1 shows a fuel tank 1. In its lower part, an inverted cylindrical guide cup 2 is fixedly fixed with perforation in the form of longitudinal cuts 3. Inside the guide cup 2 there is a metal hollow thin-walled float 4 of a cylindrical shape. The housing 5 of the microswitch is located outside the fuel tank 1. The stem 6 of the microswitch is fixed in the lower part of the guide cup 2. The sealing ring 7 provides a tight outlet of the stem 6 of the microswitch for the cavity of the fuel tank 1.

The inventive device operates as follows. In the position of the refueling fuel tank 1, the float 4 under the influence of the Archimedean force is constantly located in the upper part of the guide cup 2 and is deprived of the possibility of pressing the microswitch rod 6. During the flight, fuel is produced by a fuel intake system of any type (not considered in this application) and its level in the fuel tank 1 decreases. In this case, the float 4 moves to the lower part of the guide cup 2 under the influence of overload and without counteracting the Archimedean force and presses the microswitch rod 6 (Fig. 2) - the contacts of the microswitch are closed, thereby signaling the optimal generation of fuel from tank 1 and including mechanisms that undock the LA launch stage (for example, pyrotechnic booths Lty, castles, etc. - are not considered in this application).

The diameter of the float 4 is selected so that:

- to ensure free movement of the float 4 along the vertical axis of the guide glass 2;

- guarantee pressing the float 4 on the stem 6 of the microswitch;

- to prevent fluctuations of the float 4 inside the guide cup 2, which can lead to a turn and / or jam of the float 4 inside the guide cup 2.

The mass of the float 4 is designed to provide:

- its buoyancy;

- the necessary effort to press the float 4 on the rod 6 of the microswitch, taking into account flight overload.

The guide cup and stem of the microswitch are made of a material resistant to the aggressive environment of the fuel.

The technical result is to increase the reliability of the system for determining the moment of separation of the launch stage of a vertically launching aircraft from the LRE due to the design of the signal supply device, in which, in the process of generating fuel from the fuel tank, the float presses the microswitch rod, closing the electrical circuit.

The advantages of this device include:

- automatic response when reducing the number of intermediate control and actuators;

- reduction in the total mass of the aircraft;

- simplification of the electrical circuit of the aircraft;

- simplification and cheapening of the design of the aircraft as a whole;

- giving a command to separate the starting stage at the optimal moment: the amount of fuel in the fuel tank is sufficient to maintain the thrust of the starting engine, which ensures normal flight of the aircraft, while the amount of unused fuel is significantly reduced, since it depends only on the capabilities of the fuel intake system;

- no fuel supply is required, which will definitely remain undeveloped, which is economically viable.

The device can be made using standard equipment and materials of domestic production. Thus, the claimed invention meets the criterion of "industrial applicability".

Sources taken into account

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2. Pat. 2522537 RU, IPC ⁶ F42B 12/58, F42B 15/36. Multiple Missile / Makarovets N.A. et al. - Decl. 03/13/2013; publ. 07/20/2014, Bull. No. 20.

3. Pat. 2685161 RU, IPC ⁶ F02K 9/56. Combined fuel management system for rocket propulsion / Kurdish S.S. - Declared. 07/18/2018; publ. 04/16/2019, Bull. No. 11.

4. Pat. 2284460 RU, IPC ⁶ F42B 15/00. Rocket and drive of the separation mechanism of rocket stages / Shipunov A.G. et al. - Decl. 02/08/2005; publ. 09/27/2006, Bull. Number 27.

5. Pat. 2167388 RU, IPC F42B 15/10 (2000.01). A missile with a detachable engine / Dudka V.D. et al. - Decl. 02.24.2000; publ. 05/20/2001, Bull. No. 14.

6. Pat. 2584401 RU, IPC ⁶ F42B 15/00. The method of separation of the march stage of the aircraft and a device for its implementation / Obrazumov V.I. et al. - Decl. 02/17/2015; publ. 05/20/2016, Bull. No. 14.

7. Pat. 2874243 US, IPC G01F 23/30, G01F 23/64. Device for automatically indicating the critical oil level in engines, especially engines of motor vehicles / Metzler O. (DE) - Claim. 04/19/1957; publ. 02/17/1959.

8. Pat. 1642434 US, IPC B60K 15/03, B60K 15/06, G01F 23/30, G01F 23/64. Alarm gauge / Raye FB - Declared. 10/02/1923; publ. 09/13/1927.

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Claims (1)

A signal supply device for separating the starting stage of a vertically launching aircraft with a liquid rocket engine, characterized in that it consists of a float located with the possibility of free movement along the vertical axis inside the inverted perforated guide glass, fixedly mounted inside the fuel tank in its lower part, and microswitch, the housing of which is located outside the fuel tank, while the stem of the microswitch is placed with the possibility of free movement along the vertical axis in the lower part of the guide cup and is hermetically extended beyond the cavity of the fuel tank, and the float in its lower position is provided with the ability to press the rod of the microswitch, close its electrical contacts, signal the actual generation of fuel from the tank and turn on the separation mechanisms of the starting stage, and the force of pressing the float on the microswitch rod is provided taking into account flight overload.

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