RU2368540C1 - Hypersonic airplane and rocket propulsion system of airplane - Google Patents

Abstract

FIELD: transportation.

SUBSTANCE: airplane comprises fuselage, nose conical portion, wings, stabilisers, fuel tanks, rocket propulsion system and pipelines for supply of fuel components to it. Rocket propulsion system is installed in back part of fuselage and comprises turbopump set, two combustion chambers and flat central body in between. Rocket propulsion system of airplane comprises turbopump set with impellers of oxidant and fuel installed on shaft, turbine, start-up turbine and coaxially installed gas generator and combustion chamber. Two combustion chambers have been used, and flat central body is installed in between.

EFFECT: reduced time of speed-up.

4 cl, 4 dwg

Description

The invention relates to aviation, namely to hypersonic aircraft.

Known hypersonic aircraft according to the patent of the Russian Federation for the invention No. 20100744. The aircraft body is made in any longitudinal section along a cubic parabola with a blunt stern and a sweep angle along the leading edge of at least 60 °. The elevator is made in the form of a hinged front part of the body.

Disadvantage: relatively low flight speed of the aircraft M = 4 ... 6 ..

Known hypersonic aircraft according to the patent of the Russian Federation for invention No. 210407, containing the fuselage, the wings of the launch and mid-flight propulsion systems. Starting propulsion systems are made in the form of gas turbine engines - gas turbine engines, and marching engines - in the form of ramjets, specifically in a patented development it is proposed to use pulsating detonation air-jet engines.

The disadvantages of this aircraft: the relatively low flight speed of the aircraft and its long acceleration to hypersonic speeds due to the low thrust of the gas turbine engine.

Tasks of creating a hypersonic aircraft: reducing the time of acceleration of the aircraft to hypersonic speeds and increasing flight speed.

The solution to this problem was achieved due to the fact that a hypersonic aircraft containing the fuselage, wings, fuel tanks, marching ramjets, launch rocket engines and pipelines for supplying fuel components to them, characterized in that the launch rocket engines are made using liquid fuel components: oxidizer and fuel, an oxidizer tank and a fuel tank are installed in the fuselage, the oxidizer supply pipe passes inside the fuel tank, and the fuel tank is connected to the fuel supply pipelines with with starting rocket engines, direct-flow engines are made in the wings and are connected by fuel pipelines to the fuel tank. Starting rocket engines are made with the possibility of rotation and fixation in a vertical plane.

The proposed technical solution has novelty, inventive step and industrial applicability. The novelty is confirmed by patent research, the inventive step is that a new set of essential features made it possible to obtain a new technical effect, namely, a decrease in the acceleration time of an aircraft to hypersonic speeds and an increase in flight speed. Industrial applicability is due to the fact that the implementation of the invention does not require the creation of new unknown from the prior art parts and assemblies and new technologies.

The invention is illustrated in figure 1..3, where:

figure 1 shows a diagram of a hypersonic aircraft,

figure 2 is a section aa,

figure 3 is a rear view of the aircraft,

figure 4 is a diagram of a rocket propulsion system of an airplane.

Hypersonic aircraft (figure 1) contains the fuselage 1, the nose conical part 2, wings 3, stabilizers 4 mounted on the axles 5 with the possibility of rotation. The axis 5 is connected to the drives 6. The drive 6 is connected to the control unit 7. At the rear of the fuselage 1 there is a rocket propulsion system 8 containing two combustion chambers 9 and a central body 10. Inside the fuselage 1 there is an oxidizer tank 11 with tank insulation 12, a fuel tank 13, an oxidizer pipe 14 connected to an oxidizer tank 11 having thermal insulation 15. A fuel pipe 16 is connected to the fuel tank 16 in its lower part 16. The oxidizer tank 11 is connected to the oxidizer supply pipe 14 of a rocket propulsion system 8. The fuel tank first conduit 13 is connected to the fuel 16 with the rocket propulsion. In the front of the fuselage 1 there is a cockpit 17. In the nose conical part 2 there are four control nozzles 18 (FIG. 2) radially mounted, connected to an additional gas generator 19 and having adjustments 20. The fuselage 1 of the aircraft is mounted on the chassis 21 (FIG. 3).

The rocket propulsion system 8 (Fig. 4) contains two combustion chambers 9 and a turbopump unit (TNA) 22. The turbopump unit 22, in turn, contains an oxidizer pump impeller 24 mounted on the TNA shaft 23, a fuel pump impeller 25, a start turbine 26, additional fuel pump 27, with the shaft of the additional fuel pump 28 connected by a multiplier 29, located in the housing 30 with the shaft ТНА 23, the main turbine 31, made in the upper part of the turbopump unit 22. The gas generator 32 is mounted above the main turbine 31 coaxially with urban pump unit 22, the Combustion chamber 21 contains a nozzle 33 made of two shells and a gap "A" between them, and the head of the combustion chamber 34, inside which there is an outer plate 35 and an inner plate 36 with a cavity "B" between them. Inside the head of the combustion chamber 34, the oxidizer nozzles 37 and the fuel nozzles 38 are installed. The oxidizer nozzles 37 communicate the cavity “B” with the internal cavity of the combustion chamber “D”, and the fuel nozzles 38 communicate the cavity “B” with the internal cavity of the combustion chamber “D”. On the outer surface of the combustion chamber 21, a fuel manifold 39 is installed, from which the fuel lines 40 extend to the lower part of the nozzle 33. An outlet from the fuel valve 41 is connected to the fuel manifold 39, the inlet of which is connected by a fuel pipe 42 to the outlet of the fuel pump impeller 25. Exit from the additional the fuel pump 27 is connected by a high pressure fuel line 43 through a flow regulator 44 having an actuator 45 and a high pressure valve 46 with a gas generator 32, specifically, with a cavity “E”. The exit from the impeller of the oxidizer pump 24 through the oxidizer pipe 47 through the valve 48 is also connected to the generator 32, specifically with its cavity "G". Ignition devices 49 are installed on the head 35 of the combustion chamber 21, and ignition devices 50 are installed on the gas generator 31.

A high pressure pipe 51 is connected to the starting turbine 26 with a starting valve 52, designed to start the starting turbine 26. The other end of the high pressure pipe 52 is connected to a compressed air cylinder 53.

Electrolap devices 49 and 50, a fuel valve 41, an oxidizer valve 48, an actuator for the flow regulator 45, a high pressure valve 46, a start valve 52, and a regulator 55 installed in the gas duct 56 of one of the combustion chambers 9 are connected to the electric communication control unit 54.

A purge line 57 with a purge valve 58 is connected to the fuel manifold 39. Combustion chambers 9 can be mounted on pins 59 to swing them while controlling the course of the aircraft

Approximate characteristics of a hypersonic aircraft:

Flight speed M = 12 Starting weight, t 125 Propulsion of a rocket propulsion system, t 2 × 75 Speed gain time M = 12, sec 120

Propellant Components for LRE

Oxidizer oxygen Fuel kerosene

When starting the rocket propulsion system 8 from the control unit 7, signals are sent to the start valve 52. High pressure air from the ground system through the high pressure pipe 51 is supplied to the start turbine 26 and spins the TNA rotor 22. The oxidizer and fuel pressure at the outlet of the impellers of the oxidizer pumps 22 and fuel 23 is increasing. A signal is issued to open the valves 41, 46 and 48. The oxidizing agent and fuel enter the combustion chamber 21 and the gas generator 32. A signal is supplied to the ignition devices 49 and 50, the fuel mixture in the combustion chamber 21 and in the gas generator 32 are ignited. Starting rocket engine 4 started. The flow controller 44 using the drive 45 carry out the regulation of its mode of operation.

When you turn off the starting rocket engine from the control unit 7, a signal is supplied to the valves 41, 46 and 48 and 55, which are closed. Then, a signal is sent to open the purge valve 52, and inert gas through the purge pipe 51 enters the fuel manifold 39 and then into the cavity "A" to remove residual fuel.

At the start and acceleration of a hypersonic aircraft, the flight angles are controlled by mismatching the thrust of the combustion chambers 9 using a regulator 55, which reduces the supply of gas from the gas generator 32 to one of the combustion chambers 9. When flying in the atmosphere, the airplane is controlled by control unit 7 using actuators 6 that deflect stabilizers 4. When flying outside the atmosphere (in space) or in a rarefied atmosphere at an altitude of more than 200,000 m, an auxiliary gas generator 19 is launched (Fig. 3) and, using the regulators 20, the combustion products are fed into one from the jet nozzle control 18.

Landing of the aircraft is carried out horizontally on the chassis 21.

The application of the invention allowed the following.

1. To increase the flight speed of a hypersonic aircraft to M = 12.

2. Raise the aircraft flight ceiling to the level of space heights, as its flight and control do not need an atmosphere.

3. Accelerate the process of gaining maximum hypersonic speed.

4. Simplify the fuel supply scheme of starting and mid-flight engines.

5. Improve the reliability of hypersonic aircraft.

6. Increase the power and specific characteristics of a hypersonic aircraft,

7. Reduce the weight of the aircraft.

8. Ensure reliable control of the aircraft through the use of two control systems: aerodynamic and gas-dynamic.

9. Improve the launch and shutdown of a rocket propulsion system and ensure that they are cleaned of fuel residues after being switched off by blowing the cavities of the combustion chambers with inert gas.

Claims (4)

1. A hypersonic aircraft containing the fuselage, nose cone, wings, stabilizers, fuel tanks, rocket propulsion system and pipelines for supplying fuel components to it, characterized in that the rocket propulsion system is installed in the rear of the fuselage and contains a turbopump assembly, two combustion chambers and a flat central body between them. 2. The hypersonic aircraft according to claim 1, characterized in that four control jet nozzles are connected radially in the nose conical part and connected through regulators to an auxiliary gas generator. 3. An aircraft rocket engine installation comprising a turbopump assembly with oxidizer and fuel impellers mounted on a shaft, a turbine, a starting turbine and a coaxially mounted gas generator, connected by gas ducts to combustion chambers, characterized in that two combustion chambers are used and a flat central body is installed between them. . 4. The rocket propulsion system of an aircraft according to claim 3, characterized in that a regulator is installed in the gas duct of one of the combustion chambers.

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