RU2605797C1 - Aircraft hydraulic system - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: invention relates to aircraft engineering and can be used in hydraulic systems designs implementing several control modes. Hydraulic system comprises electrically driven pump (7) with controlled supply, actuator motor (8), presented in form of hydraulic cylinders with pistons progressive motion, main line (10), connecting actuating motor (8) with electrically driven pump (7). System is equipped with hydraulic tank (1), inside of which piston with spring-loaded dampers with possibility of longitudinal movement is installed, which separates hydraulic tank (1) to high-pressure and low-pressure chambers. Power-driven pump (7) input is connected to hydraulic tank low-pressure chamber, and its output is connected with hydraulic tank high-pressure chamber and to actuator motor (8) input, wherein electrically driven pump (7) and actuator motor (8) are connected to missile control system (12).

EFFECT: reduction of total heat release, dimensions and weight by required increase of hydraulic system power during aircraft with hydraulic tank programmed manoeuvres periods.

1 cl, 2 dwg

Description

The invention relates to rocket and aeronautical engineering and can find primary application in the design of hydraulic systems that implement several control modes, for example, in hypersonic missiles.

A well-known hydraulic system (see DP Popov "Mechanics of hydraulic and pneumatic drives", publishing house of MSTU named after Bauman, 2002, p. 18, Fig. 1.2), which has a pump with variable flow rate, an executive motor, presented in the form hydraulic cylinders with translational movements of the pistons, a line connecting the executive engine with an electric drive pump. This system implements a single-mode type of control. However, for the implementation of several flight control modes that differ in power requirements several times at the beginning of the flight and on the marching section of the trajectory in dense layers of the atmosphere, this system is not practical. Due to the fact that the volume of the hydraulic accumulator is limited, in order to increase the power of the hydraulic system, the available power of the hydraulic pump has to be increased several times, which leads to an increase in the mass and dimensions of the hydraulic system, and also increases the heat generation and volume of the working fluid.

The aim of the present invention is to reduce the total heat, dimensions and mass by the required increase in the power of the hydraulic system during periods of program maneuvers of an aircraft with a hydraulic tank.

This goal is achieved by the fact that the hydraulic system of the aircraft contains an electric drive pump with adjustable flow, an executive motor, presented in the form of hydraulic cylinders with translational movements of the pistons, a line connecting the executive motor to the electric drive pump. The hydraulic system is equipped with a hydraulic tank, inside of which a piston with spring-loaded dampers is mounted with the possibility of longitudinal movement, dividing the hydraulic tank into high-pressure and low-pressure cavities. The input of the electric drive pump is connected to the low-pressure cavity of the hydraulic tank, and its output is connected to the high-pressure cavity of the hydraulic tank and to the entrance to the executive engine (steering units), while the electric drive pump and the executive engine are connected to the rocket control system (SU).

The invention is illustrated by drawings, where in FIG. 1 shows a diagram of the hydraulic system of an aircraft; FIG. Figure 2 shows a fragment of the system during operation: on the left is a fragment of the hydraulic system in the steady-state first mode of operation, on the right is a fragment of the hydraulic system in the second mode - the hydraulic accumulator mode due to the displaced working fluid from the high-pressure cavity.

In FIG. 1-4, the positions are indicated in the following order:

1 - hydraulic tank;

2 - differential piston;

3 - damper;

4 - high-pressure cavity;

5 - low pressure cavity;

6 - pipeline;

7 - electric drive hydraulic pump;

8 - drive electric motor;

9 - check valve;

10 - pressure line;

11 - steering units;

12 - control system;

13 - electric cable.

The hydraulic system of the aircraft contains a hydraulic tank 1, divided by a differential piston 2 with spring-loaded dampers 3 into a high-pressure 4 and a low-pressure 5 cavity, connected by a pipe 6 to a dual-mode electric drive hydraulic pump 7. The output of the hydraulic pump 7 is connected through a non-return valve 9 to a pressure line 10 and then to the steering units 11. At the same time, the hydraulic pump 7 and steering units 11 are connected to the control system 12 of the aircraft using electric cables 13.

The operation of the hydraulic system is as follows.

In the first operating mode, the electric drive hydraulic pump 7 is started. The working fluid through the check valve 9 through the pressure line 10 enters the steering units 11 and at the same time into the high-pressure cavity 4 of the hydraulic tank 1, creating through the piston 2 the pressure necessary for normal cavitation-free operation of the hydraulic pump 7. the cavities of the spring-loaded dampers 3 located in the differential piston 2 are filled with a working fluid and the hydraulic system operates in the minimum system power mode corresponding to the flow from leaks through the steering units 11.

To ensure stabilization of the aircraft, the hydraulic system operates in the second mode with maximum power. The control system 12 through electric cables 13 issues a command for supplying a working fluid through a non-return valve 9 through a pipe 6 to the high-pressure cavity 4 of the hydraulic tank 1 and to the steering units 11, which at that time take up large aerodynamic loads. When this check valve 9 closes the outlet of the hydraulic pump 7, reducing the power consumed by the electric motor 8.

At the end of operation in the second mode, the hydraulic pump 7 gradually stops and the flow of the working fluid flows to the steering units 11 during operation of the hydraulic system with the consumption of the usual power mode. If a peak flow rate occurs in the hydraulic system (in any modes), the pressure in the pressure line 10 decreases and under the action of a pressure drop determined by the ratio of the areas, the spring-loaded dampers 3 are discharged. The differential piston 2 moves by a value A, displacing, as a hydraulic accumulator, the missing volume of the working fluid.

Thus, the proposed technical solution made it possible to implement a hydraulic system with a reduced total heat release and to reduce the dimensions and weight by the required increase in the hydraulic system power during periods of program maneuvers of the aircraft.

Claims (1)

The hydraulic system of the aircraft, containing an electric drive pump with adjustable flow, an executive engine, presented in the form of hydraulic cylinders with translational movements of the pistons, a line connecting the executive engine with an electric drive pump, characterized in that the system is equipped with a hydraulic tank, inside of which the piston is mounted with the possibility of longitudinal movement with spring-loaded dampers, dividing the hydraulic tank into high-pressure and low-pressure cavities, electric input the drive pump is connected to the low-pressure cavity of the hydraulic tank, and its output is connected to the high-pressure cavity of the hydraulic tank and to the entrance to the executive engine, while the electric drive pump and the executive engine are connected to the rocket control system.

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