RU2378166C1 - Multi-stage rocket carrier, method of its launching and nuclear rocket engine - Google Patents

Abstract

FIELD: aerospace engineering.

SUBSTANCE: invention relates to aerospace engineering, namely, to liquid propellant rocket engines running on, at a time, nuclear and rocket fuel, i.e. oxidiser and fuel. Proposed carrier rocket comprises rocket pods of the 1^st and 2^nd stages connected in parallel and furnished with fuel and oxidiser tanks. 2^nd stage pod incorporates nuclear reactor. 1^st and 2^nd stage engines comprises combustion chamber and turbopump unit. Said combustion chamber accommodates heat exchangers communicated via circulation lines with nuclear reactor. 2^nd stage engine incorporates nozzle extension moving along rocket axis. Proposed method comprises starting 1^st and 2^nd stage engines. Note here that nuclear reactor and heat carrier circulation are initiated simultaneously with starting of engines. Combustion products heated sufficiently, oxidiser flow rate is reduced to the level required to drive turbopump unit. 1^st stage engines out, heat carrier circulation through 1^st stage heat exchangers is cut off, and 1^st stage rocket pods are separated. After said separation, nozzle extension is moved out for one or several engines of 2^nd stage to increased nozzle expansion degree. Rocket nuclear engine comprises combustion chamber with regenerative cooling system and turbopump unit. The latter comprises oxidiser and fuel pumps, primary and starting turbines and gas generator. Fuel pump outlet communicates with regenerative cooling system inlet, while regenerative cooling system outlet communicates with gas generator inlet. Oxidiser pump outlet communicates with gas generator second inlet. Aforesaid combustion chamber houses heat exchanger communicated via heat carrier circulation lines with nuclear reactor.

EFFECT: perfected performances in wide range of flight characteristics at various altitudes.

3 cl, 4 dwg

Description

The invention relates to rocket technology, specifically to multi-stage rockets with liquid rocket engines.

Known multi-stage launch vehicle according to the patent of the Russian Federation for the invention No. 2137680, publ. September 20, 1999. It contains several rocket stages with engines installed in series. Engines on rocket stages start sequentially. The disadvantage is the low power ratio of the rocket at launch.

Known multi-stage launch vehicle, adopted as a prototype, according to the patent of the Russian Federation for invention No. 2116941, publ. 08/10/1998 g, IPC B64G 1/00. It contains parallel-connected rocket blocks of the first and second stages of the launch vehicle, connected by power communication units and equipped with engines of the first and second stages. At the start, the engines of the first and second stages are turned on simultaneously, at a certain height (practically in a vacuum), a command is sent from the control unit to disconnect the power communication nodes and reset the rocket blocks of the first stage, while the engine (engines) of the second stage continues to work.

Liquid propellant rocket engines are often used as marching engines for high-powered launch vehicles; they are easier to regulate compared to solid-fuel ones.

Known liquid rocket engine according to the patent of the Russian Federation for invention No. 2095607, intended for use as part of space booster blocks, stages of rocket launchers and as a main engine of spacecraft including a combustion chamber with a regenerative cooling path, pumps for supplying components - fuel and an oxidizer with a turbine on one the shaft into which the capacitor is inserted. The condenser outlet through the refrigerant line is connected to the entrance to the combustion chamber and to the entrance to the regenerative cooling path of the combustion chamber. The exit from the condenser through the coolant is connected to the inlet to the pump of one of the components. The output from the pump of the same component is communicated with the condenser inlet through the refrigerant line. The second input of the condenser is in communication with the output of the turbine. The output of the pump of the other component is communicated with the entrance to the combustion chamber.

The disadvantage of the engine according to the patent of the Russian Federation No. 2095607 is the deterioration of the cavitation properties of the pump during condensate bypass.

A known method of operation of the liquid propellant rocket engine and a liquid rocket engine according to the patent of the Russian Federation for the invention No. 2187684. The method of operation of a liquid-propellant rocket engine is to supply fuel components to the combustion chamber of the engine, gasify one of the components in the cooling path of the combustion chamber, supply it to the turbine of the turbopump unit, and then discharge it into the nozzle head of the combustion chamber. Part of the flow rate of one of the fuel components is directed to the combustion chamber, and the remaining part is gasified and directed to turbines of turbopump units. The gaseous component spent on the turbines is mixed with the liquid component entering the engine at a pressure higher than the saturated vapor pressure of the resulting mixture. A liquid propellant rocket engine comprises a combustion chamber with a regenerative cooling path, fuel component supply pumps, and a turbine. The engine comprises a booster turbopump pump and a mixer installed in series in front of the feed pump of one of the fuel components of the main turbopump assembly. The pump outlet of the main turbopump assembly is connected both to the nozzle head of the combustion chamber and to the regenerative cooling path of the combustion chamber. The regenerative cooling circuit, in turn, is connected with the turbines of the main and booster turbopump units, the outputs of which are connected to the mixer.

The disadvantage of this scheme is that the thermal energy removed during cooling of the combustion chamber may not be enough to drive a turbopump engine unit of very high power.

Known LRE according to the patent of the Russian Federation for the invention No. 2190114, IPC 7 F02K 9/48, publ. 09/27/2002 This LPRE includes a combustion chamber with a regenerative cooling path, a TNA turbopump unit with oxidizer and fuel pumps, the output lines of which are connected to the head of the combustion chamber, the main turbine and the drive circuit of the main turbine. The main turbine drive circuit includes a fuel pump and a regenerative cooling path of the combustion chamber connected in series with each other and connected to the main turbine inlet. The exit from the turbine TNA is connected to the input of the second stage of the fuel pump.

This engine has a significant drawback. Bypassing the fuel combustion chamber heated in the regenerative cooling path to the entrance to the second stage of the fuel pump will lead to cavitation. Most LREs use fuel components such that the oxidizer consumption is almost always greater than the fuel consumption. Therefore, for powerful rocket engines with great thrust and high pressure in the combustion chamber, this scheme is unacceptable, because fuel consumption will not be enough to cool the combustion chamber and drive the main turbine.

In addition, the LRE launch system, the ignition system of the fuel components and the LRE shutdown system and its cleaning of fuel residues in the regenerative cooling path of the combustion chamber have not been developed.

Known liquid rocket engine according to the patent of the Russian Federation for the invention No. 2232915, publ. September 10, 2003, which contains a chamber, a turbopump unit, a gas generator, a launch system, means for igniting fuel components and fuel lines. The output of the oxidizer pump is connected to the inlet of the gas generator. The output of the first stage of the fuel pump is connected to the channels of regenerative cooling of the chamber and to the mixing head. The output of the second stage of the fuel pump is connected to an electric flow regulator. Another input of the regulator is connected to the starting tank with standard fuel. The output from the regulator is connected to a gas generator. The outlet of the gas generator is connected to the inlet of the turbine pump unit, the outlet of which is connected to the mixing head. The flow regulator is equipped with a hydraulic actuator of the preliminary stage, which is connected through the cavitating nozzle and hydraulic relay to the starting tank with standard fuel. The hydraulic relay is connected to the second stage of the fuel pump. The throttle installed at the output of the first stage of the fuel pump is made in conjunction with a controlled valve of the preliminary stage.

The main element of the output device of a jet engine and a combined thrust engine is a jet nozzle. In the jet nozzle, gas expands from the turbine or afterburner of the gas turbine engine or from the combustion chamber (or other device for heating the working fluid) of the rocket engine, accompanied by an increase in its speed and kinetic energy. The gas expansion in the jet nozzle occurs to the ambient pressure at the design mode of the nozzle and to a pressure different from the ambient pressure at off-design nozzle modes. The rate of gas outflow from the jet nozzle of the jet engine in the nozzle design mode determines at a given flight speed the specific thrust of the engine. The rate of gas outflow from the jet nozzle of a rocket engine in the nozzle design mode determines the specific thrust of the engine, regardless of flight speed. The speed of gas outflow from the jet nozzle of modern jet engines under terrestrial static conditions reaches 1000 m / s and more for air-jet engines and up to 3000 m / s and more for rocket engines. Distinguish between adjustable and unregulated jet nozzle. The adjustable nozzle is equipped with a device for changing its cross section during engine operation. In a subsonic tapering nozzle, the regulation consists, as a rule, in changing the area of the exit section of the nozzle. In a supersonic nozzle, both the critical section area and the nozzle exit section area are subject to regulation. The adjustable nozzle is used in turbojet engines with an afterburner, as well as in some other gas turbine, jet and rocket engines. The nozzle of a rocket engine is also called the nozzle of the engine chamber, or simply a nozzle and its regulation is practically not applied due to the very high temperature of the expiration of the combustion products.

The task of creating the invention: ensuring the optimal operation of rocket engines of the second stage in a wide range of flight modes at different heights, simplifying the pneumohydraulic circuit, increasing reliability, increasing the power and technical characteristics of the rocket engine.

The task of creating the invention: improving the technical characteristics of the rocket in a wide range of flight modes at various heights.

The solution of these problems was achieved in a multi-stage launch vehicle containing parallel-mounted rocket blocks of the first and second stages of the launch vehicle with oxidizer and fuel tanks, connected by power communication units and equipped with engines of the first and second stages, characterized in that in the block of the second stage a nuclear reactor is installed, the engines of the first and second stages are made containing a combustion chamber and a turbopump unit, heat exchangers connected to pipelines are installed inside the combustion chambers and recirculation with a nuclear reactor, the second stage engine is made with a nozzle nozzle having the ability to move along the axis of the rocket.

The solution of these problems was achieved in the method of launching a multi-stage launch vehicle containing rocket blocks of the first and second stages, including starting the engines of the first and second stages, turning off the engines of the first stage and resetting the blocks of the first stage, characterized in that at the same time as starting the engines, a nuclear reactor and coolant circulation, simultaneously with turning off the first stage engines, turn off the coolant circulation through the heat exchangers of the first stage engines, reset the rocket blocks of the first stage, after dropping the rocket blocks of the first stage at each engine of the second stage, the nozzle nozzle is extended to the lower position to increase the degree of expansion of the nozzle.

The solution of these problems was achieved in a nuclear rocket engine containing a combustion chamber with a regenerative cooling system and a turbopump unit, in turn containing oxidizer and fuel pumps and a main and starting turbine, a gas generator installed coaxially with the turbopump unit, characterized in that the fuel pump exits connected to the entrance to the regenerative cooling system, the output of which is connected to the entrance to the gas generator, the output from the oxidizer pump is connected to the second entrance to the gas generator, inside the chamber ry combustion installed heat exchanger connected with the coolant recirculation piping of a nuclear reactor.

The invention is illustrated in figure 1 ... 4, where:

figure 1 shows a diagram of a launch vehicle,

figure 2 shows a diagram of a nuclear rocket engine,

figure 3 shows a view for a nozzle nozzle made of carbon-carbon composite material,

figure 4 shows a view And for a nozzle nozzle made cooled.

The launcher contains at least one rocket block of the first stage 1 with engines of the first stage 2 having a combustion chamber 3 and a rocket block of the second stage 4 with engines of the second stage 5 having a combustion chamber 6 and nozzle nozzles 7. To nozzle nozzles 7 drives 8 are connected. Missile blocks of the first and second stages 1 and 4 are connected by power communication units 9. A control unit 10 is mounted on the rocket, connected by electrical connections 11 to the engines of the first and second stages, respectively 2 and 5, and with power communication nodes 9. Inside the rockets of the first stage 1 unit, an oxidizer tank 12 and a fuel tank 13 are installed. Inside the second stage 4 rocket unit, an oxidizer tank 14 and a fuel tank 15 are installed, as well as a nuclear reactor 16, which has coolant circulation pipelines 17 and 18, in one of which a circulation pump is installed 19 is connected to a heat exchanger (heat exchangers) 20 installed in the combustion chamber 3 of the first stage engine 2, and a heat exchanger (heat exchangers) 21 installed in the combustion chamber 6 of the second stage engine 5.

The second stage 5 liquid propellant rocket engine contains a combustion chamber 6 and a turbopump unit 22. The TNA-22 turbopump unit (Fig. 2), in turn, contains (Fig. 2) an oxidizer pump 23, a fuel pump 24, an additional fuel pump 25, and a turbine 26 and a gas generator 27.

The gas generator 27 is mounted above the turbine 26 coaxially with the turbopump unit 22 (figure 2). The combustion chamber 5 contains a nozzle 28 made of two shells 29 and 30 with a gap "B" between them, and the head of the combustion chamber 31. Inside the head of the combustion chamber 31, oxidizer nozzles 32 and fuel nozzles 33 are installed. A collector is installed on the outer surface of the combustion chamber 5. fuel 34. To the fuel manifold 34 is connected a high pressure pipe 35, coming from the additional fuel pump 25. The input of the additional fuel pump 25 is connected by a pipe 36, in which the flow regulator 37 with the drive 38 is installed, with the output of the hot pump 24. The second output of which the fuel pump 24, a conduit 39, which establishes the flow regulator 40 with a drive 41 and fuel valve 42, is connected to the gas generator 27. The output of the pump 23 the oxidizer oxidant conduit 43 through the oxidant valve 44 is also connected to the gas generator 27.

The engine has a control unit 10, to which the fuel valve 42, the oxidizer valve 44, the drive 41 of the flow regulator 40 are connected by electrical connections 45. Each engine of the second stage 5 (Fig.2 ... 4) contains a nozzle nozzle 7, made along the profile as a continuation of the nozzle 28 and having the ability to move along the axis of the combustion chamber 6 using the actuator 8. The actuator 8 can be made of one or more actuators 46. The actuator 46 is connected on one side by a rod 47 to a power plate 48, and on the other hand to a camera with Gorania 6. It is preferable to use three actuators 46 mounted circumferentially symmetrically to the axis of the combustion chamber 6 and connect them with a synchronization mechanism, for example, a belt one. For synchronization, an electrical or mechanical device, for example with a chain or with a belt, can be used. Electrical synchronization can be implemented using 46 electric motors as actuators.

Figure 3 and 4 shows the design of the junction of the nozzle nozzle 7 and the nozzle 28, while figure 3 shows the nozzle nozzle 7 made of graphite-graphite composite material, only the upper ring 49 is made of metal, it can be combined with the power plate 48 The docking nozzle nozzle 7 with the lower part of the nozzle 28 is made to ensure tightness on the conical surface "B". Figure 4 shows a cooled nozzle nozzle 7, which contains a cooling jacket 50, forming a gap “B” with the nozzle nozzle 7, the cavity between the nozzle nozzle 7 and the nozzle 28 is connected by flexible pipelines 51 and 52, respectively, with the exit of the fuel pump 24 and with the collector fuel 34 (not shown in FIGS. 3 and 4).

To prevent leakage of the coolant during undocking of the first stage of the launch vehicle, check valves 53 (FIG. 2) are installed in the circulation lines 17 and 18.

When starting the engines of the first and second stages, respectively 3 and 5, from the control unit 10, signals are given to start the TNA 22. The pressure of the oxidizer and fuel at the outlet of the oxidizer pumps 23, the fuel pump 24 and the additional fuel pump 25 increases. A signal is issued to open the valves 42 and 44. The oxidizing agent and fuel enter the combustion chamber 6 and the gas generator 27. A signal is supplied to the ignition devices (not shown in FIGS. 1 ... 4), and the fuel mixture in the combustion chamber 6 and in the gas generator 27 is ignited. Engines of the first and second stages, respectively 3 and 5 are started. The flow regulator 40 for each engine regulates the operation mode by changing the ratio of fuel components in the gas generator 27. The engine of the second stage 5 operates on the chemical components of the fuel. After that, the nuclear reactor 16 is started up or advance, and the coolant is fed through the circulation pipe 17 to the heat exchangers 20 and 21. Using the pump 19, the combustion products in the combustion chambers 3 and 6 are additionally heated, and the engines 2 and 5 will create a large traction force, while the oxidizing agent is reduced to the level necessary to drive the TNA 22.

The launch vehicle can be controlled in one of three ways:

1. The mismatch of the traction force of the opposing engines.

2. The use of steering engines or combustion chambers.

3. By swinging the marching engines of both the first and second stages.

After the rocket has climbed the altitude, the control unit 10 gives a signal to the power communication nodes 9, they break and the missile blocks of the first stage 1 are thrown to the sides. Then, a signal is sent from the control unit 10 to the actuator (s) 47, which moves the nozzle nozzle 7 (nozzle nozzles) to the lowermost position. The length of the nozzle (s) and the degree of expansion of the combustion products in it (in them) increases. The combustion products flowing from the nozzle 28 are further expanded in the nozzle (s) 7 to the ambient pressure, creating additional traction without increasing fuel consumption and increasing the power of a nuclear reactor. This leads to an improvement in the specific technical characteristics of LRE at high altitudes, primarily to a decrease in specific fuel consumption (fuel consumption per unit of thrust). The use of the synchronization mechanism and sealing the junction of the nozzle with the nozzle nozzle on the conical surface "B" eliminates the bias and the formation of a gap between them, leading to leakage of combustion products into this gap with a deterioration in engine performance. If the engine of the second stage 5 is designed with a high degree of expansion, without adjusting it in height, this will lead to the fact that at high altitudes the flight of the launch vehicle, the engine will operate in optimal mode, and at low altitudes its technical characteristics will deteriorate sharply due to overexpansion of gases inside the nozzle. In addition, an uncontrolled lateral component of the nozzle thrust force arises and burnout of the nozzle inner wall is possible at the points of separation of the supersonic flow.

When the engine is turned off, a signal is sent from the control unit 10 to shut down the nuclear reactor 16 and to the valves 42 and 44, which are closed.

The application of the invention allowed:

1. Significantly increase the flight range of the rocket with its identical starting weight due to the use of nuclear fuel.

2. To provide good technical characteristics of the second-stage engine in a wide range of modes of its operation at various heights.

3. To ensure reliable operation of the nozzle nozzle at high temperatures.

4. Eliminate skew when extending the nozzle nozzle to the lower position due to the use of the synchronization mechanism.

5. To ensure the tightness of the junction of the nozzle nozzle with the nozzle for both versions of the nozzle nozzle: uncooled and cooled.

Claims (3)

1. A multi-stage launch vehicle containing parallel-mounted rocket blocks of the first and second stages of the launch vehicle with oxidizer and fuel tanks, connected by power communication units and equipped with engines of the first and second stages, characterized in that a nuclear reactor is installed in the second stage unit , the engines of the first and second stages are made containing a combustion chamber and a turbopump assembly, heat exchangers are installed inside the combustion chambers, connected by recirculation pipelines to a nuclear reactor, the engine of the second stage is made with a nozzle nozzle having the ability to move along the axis of the rocket. 2. A method of launching a multi-stage launch vehicle containing rocket blocks of the first and second stages, comprising starting the engines of the first and second stages, turning off the engines of the first stage and resetting the blocks of the first stage, characterized in that at the same time as the engines are started, they include a nuclear reactor and coolant circulation, at the same time as the first stage engines are turned off, the coolant circulation is switched off through the heat exchangers of the first stage engines, the first stage missile blocks are dropped, after the reset rocket blocks of the first stage at each engine of the second stage extend the nozzle of the nozzle in the lower position to increase the degree of expansion of the nozzle. 3. A nuclear rocket engine containing a combustion chamber with a regenerative cooling system and a turbopump assembly, in turn, comprising oxidizer and fuel pumps, a main and starting turbine, a gas generator mounted coaxially with the turbopump assembly, characterized in that the exit from the fuel pump is connected to the entrance to the regenerative cooling system, the output of which is connected to the entrance to the gas generator, the exit from the oxidizer pump is connected to the second entrance to the gas generator, a heat exchange is installed inside the combustion chamber to connected piping recirculating coolant to the nuclear reactor.

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