RU2514570C1 - Device for regenerative cooling of liquid-propellant engine supersonic section - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to aerospace engineering and can be used for cooling the liquid-propellant rocket engine nozzle supersonic section. Invention aims at developing a serviceable nozzle supersonic section cooler with lower coolant pressure (Pcool<<Pc) to produce high efficiency liquid-propellant rocket engines with increased chamber pressure, simplifying the nozzle design and increasing their efficiency. Target is hit by incorporating a check valve in heat carrier circulation line at main line communicating nozzle supersonic section cooling circuit outlet with turbine inlet and connecting heat carrier tank with valve with said main line at section between check valve and turbine. Besides, exhaust pipe with valve or receiver is connected at main line section between nozzle supersonic section and check valve.

EFFECT: higher reliability and efficiency, lower costs.

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Description

The invention relates to rocket and space technology and can be used to cool the supersonic part of the nozzle of liquid rocket engines (LRE).

Years of experience in creating domestic marching rocket engines for the first stages of launch vehicles (LV) show that the main way to increase their efficiency is to increase the pressure in the chamber. When the chamber is cooled by a fuel component, which is then burned in it, the pressure in the cooling path (Rohl) is always greater than the pressure in the chamber (Pk). This is due to the fact that the liquid propellant rocket engine is a continuous engine (unlike, for example, cyclically operating internal combustion engines), therefore, to ensure a continuous flow of the cooler in the area from the entrance to the cooling path to the firing cavity of the Rokhl chamber, it must constantly exceed Pk by the amount of hydraulic resistances of this section (ΔР of the cooling path + ΔР of the mixing head + ΔР of regulators ... etc.). The pressure level in the chamber of modern rocket engines reaches 30 MPa, while the pressure in the cooling path reaches 45 MPa. In the patents of the Russian Federation [1,2] it is indicated that a further increase in pressure in the cooling path is dangerous, since it can destroy the mechanical bonds between the inner and outer shells of the chamber. This problem is especially acute for engines with afterburning of reducing gas, the chamber of which is cooled by fuel.

In order to reduce the pressure in the cooling path without reducing Pc, one can refuse to cool the chamber with the components of the fuel burned in it. Such a path is proposed in the patent of the Russian Federation [3] (analogue). In the proposed device, the cooling system of the chamber, in which the coolant, which is not a component of the fuel burned in the chamber, circulates in a closed circuit containing a pump, a regenerative cooling path of the combustion chamber (KS), a turbine, a pump and a heat exchanger.

In the path of regenerative cooling of the compressor, the chamber wall is cooled, and the coolant is heated, turning into steam. Next, the resulting steam is triggered on a turbine driving the pump. After the turbine, the coolant is supplied to the heat exchanger, where it is cooled or condensed due to heat exchange with the coolant. Then, the coolant pressure is pumped up in the pump and it again enters the inlet to the regenerative cooling circuit of the compressor station. The proposed cooling system allows, through the use of an intermediate coolant and its increased pressure drops in the regenerative cooling path of the compressor, to intensify the heat sink from the internal wall of the compressor rocket engine and increase the pressure level in the compressor. This leads to an increase in engine specific impulse and increases fuel density. However, the implementation of such a cooling scheme is associated with significant difficulties. When cooling the entire chamber in order to ensure an acceptable thermal state of the critical section area, a large flow rate of the coolant is required and, accordingly, problems arise with the mass-dimensional characteristics of the heat exchanger. In addition, the high hydraulic resistance of the cooling duct in the region of the critical section limits the possibilities of linking the power balance between the pump and the coolant turbine. Therefore, in the practical implementation of the proposed camera cooling system, these drawbacks may exceed the obtained beneficial effect.

At the same time, the level of risk of breaking bonds between the shells of the chamber is significantly different for its different sections. In the region of high heat fluxes (i.e., on the chamber block including the combustion chamber, the nozzle inlet and the critical section area), the number of ribs and, accordingly, the number of mechanical bonds between shells created by soldering along the vertices of these ribs is so large which is redundant in terms of ensuring the strength of the bonds between the shells. And this is due to the fact that the number of ribs and their geometric characteristics are selected primarily from the conditions for ensuring reliable cooling of the chamber (the maximum amount of heat transferred by the side surfaces of the ribs to the cooler, ensuring the required speed of movement of the cooler), and the necessary bond strength is automatically ensured between the shells. So, according to the results of computational and experimental studies, the level of pressure that destroys these bonds exceeds the working pressure in the cooling path several times. Thus, the pressure in the cooling path of the combustion chamber unit is not a factor limiting the increase in Pk. Therefore, the cooling of the camera unit may well be carried out in the traditional way, i.e. high pressure fuel component.

The situation is different in the supersonic part of the nozzle. Heat fluxes in the supersonic part of the nozzle are significantly less than in the chamber unit. This leads to the fact that much less stringent requirements are imposed on the cooling system and therefore the number of ribs is determined mainly only by considerations of sufficient strength of the bonds between the shells and sufficient rigidity of the fire wall between the ribs. However, at high pressures in the cooling path it is very difficult to provide large reserves of strength of the supersonic part of the nozzle. This is due to significant overall dimensions compared to the camera unit. So, for example, if the cooling path of the supersonic part of the nozzle is performed with a fin step close to the step of the ribs in the critical section, the nozzle will turn out to be unacceptably heavy. You can also note the complexity and complexity of manufacturing soldered nozzles, the presence of poorly predicted loads (for example, vibrational) - all this increases the risk of destruction of the nozzles, especially under conditions of ultrahigh pressure in the cooling path.

In this regard, when creating modern high-pressure rocket engines with a high pressure in the chamber, especially for the lower stages of the launch vehicle, it is advisable to reduce the pressure in the cooling path not in the entire chamber (as is done in the analogue [3]), but only in the least durable part of it - in the supersonic part of the nozzle.

In [4] (prototype), a scheme of an oxygen-methane liquid propellant rocket engine with afterburning of a reducing generator gas was proposed, in which the chamber unit is cooled by high-pressure fuel, which then goes to the gas generator, turbine, and the mixing head of the chamber, and the supersonic part of the nozzle is cooled by low-pressure ammonia, circulating in a closed circuit, and has its own, independent of the fuel component supply system, turbopump unit. Ammonia vapor after cooling the nozzle rotates the turbine, then liquefies in a heat exchanger (refrigerant is the flow of liquid oxygen burned in the chamber) and is compressed by the pump, after which it again enters the nozzle cooling path. The pump is driven by a turbine rotated by ammonia vapor preheated after the cooling path. In this case, the pressure in the ammonia circuit is significantly lower, even than the pressure in the chamber. The proposed method for implementing the above idea of reducing the pressure of the cooler not in the entire cooling path of the chamber, but only in the supersonic part of the nozzle has the following disadvantages:

- the scope of the proposed cooling scheme is unreasonably narrowed (oxygen-methane LRE with afterburning of the reducing generator gas) - such a scheme can be used for any high-pressure LRE in the chamber;

- the choice of ammonia as a coolant for the cooling circuit, and liquid oxygen as a coolant can lead to freezing of ammonia and blockage of the heat exchanger.

In addition, from the point of view of the thermal state, the proposed scheme works only in stationary mode. At the time of starting the engine, ammonia is not heated, the charge pump does not rotate, therefore there is no circulation of the coolant along the cooling circuit of the supersonic part of the nozzle. Accordingly, there is no cooling, which almost instantly leads to overheating of the wall and to burnout.

The objective of the invention is to create a transient and stationary operation device for cooling the supersonic part of the nozzle with a low level of cooler pressure (Rokhl << Rk), which should provide the ability to create highly economical liquid-propellant rocket engines with increased pressure in the chamber, while simplifying the manufacture of nozzles and increasing them reliability.

The solution to this problem is achieved by the fact that in the device for regenerative cooling of the supersonic part of the LRE nozzle, which includes a heat carrier tank with a valve and a coolant circulation circuit, consisting of a cooling path of the supersonic part of the nozzle, turbine, heat exchanger, pump and mains connecting them, in the heat carrier circulation circuit to a check valve is installed on the line connecting the output of the cooling path of the supersonic part of the nozzle to the turbine inlet, and a coolant tank with a valve is connected to this line at the section between the non-return valve and the turbine, in addition, at the section of the main line between the output of the cooling path of the supersonic part of the nozzle and the non-return valve, an exhaust pipe with a valve or receiver is connected.

The presence of a non-return valve and the introduction of an exhaust pipe with a valve or a receiver into the coolant circuit allow reliable operation of the chamber in stationary mode with a small flow rate of the coolant that circulates in a closed circuit that includes a coolant pump, a supersonic nozzle cooling path, a check valve, a turbine, the heat transfer pump and the heat exchanger, which rotate. In the engine start mode, if a supersonic part of the nozzle is connected between the outlet of the cooling path of the nozzle and the exhaust pipe check valve with the valve, the coolant flows through an open circuit, including a starting tank with a coolant reserve, a valve, a turbine, a heat exchanger, a coolant pump, and a supersonic part cooling path nozzles and exhaust pipe with a valve installed on it. When the engine starts, when connected to this section of the receiver line, the coolant circulates from the very beginning in a closed loop. It is obvious that the exhaust pipe has a lower mass than the receiver, however, when using the pipe, the coolant is consumed, which may block its weight advantages. Apparently, the feasibility of using an exhaust pipe with a valve or receiver will be determined primarily by the engine start sequence diagram. So, for engines with a “cannon” start, it is preferable to more energetically cool the nozzle during start-up, i.e. the use of an exhaust pipe, and for engines that start quite smoothly through an intermediate stage of traction, it is preferable to use a receiver.

The low pressure in the coolant circulation circuit is primarily ensured by the fact that there is no need to supply cooler to the mixing head of the chamber and, in addition, the cooling of the supersonic part of the nozzle does not require high flow rates of the cooler and, accordingly, high pressure in the cooling path to compensate for hydraulic losses. When using the proposed regenerative cooling device, for example, as part of a first-stage engine with Pk = 17 MPa, the maximum pressure of the coolant will not exceed 5 MPa. This will allow you to create a rocket engine with high pressure in the chamber, but without increasing the risk of destruction of the supersonic part of the nozzle. In addition, this level of pressure of the cooler will allow you to abandon the brazed-milled design of the cooling path of the supersonic part of the nozzle and return to the design in which the inner and outer shells are connected by spot welding in places of local stampings in the outer shell. This will remove the currently existing restrictions imposed on the dimensions of the nozzle by the dimensions used in the industry of vacuum furnaces, and significantly reduce the complexity of manufacturing nozzles.

The proposed device for regenerative cooling of the supersonic part of the LPRE nozzle is illustrated by the presented schemes of Fig.1 and Fig.2. Figure 1 shows a device for regenerative cooling of the supersonic part of the nozzle of the rocket engine with a connection on the section of the line between the output of the cooling path of the supersonic part of the nozzle and the check valve of the exhaust pipe equipped with a valve. Figure 2 is a device with a connection on this section of the receiver line.

A device for regenerative cooling of the supersonic part of the LPRE nozzle (FIG. 1, FIG. 2) includes a heat carrier tank 1 with a valve 2, a turbine 3, a heat exchanger 4, pump 5, a cooling path for the supersonic part of the nozzle 6, a common turbine shaft and pump 7, check valve 8, valve 9 of the exhaust pipe and receiver 10.

The proposed device (Figure 1) operates as follows. In the tank 1 under pressure contains the coolant. After the engine starts, valves 2, 9 open and the coolant starts to flow from the coolant tank 1 to the line between the check valve 8 and the turbine 3. At the same time, the check valve 8 is closed, so the coolant moves in the direction of the turbine 3, enters the blades of its impeller and starts to spin turbine 3, pump 5 starts working through the common shaft 7. After the turbine, the coolant passes through the heat exchanger, the pump enters the cooling path of the supersonic part of the nozzle 6 and cools the nozzle wall starting to warm up, then the coolant is discharged through the exhaust pipe with an open valve 9 into the atmosphere or into the exhaust stream of the engine. The movement of the coolant at the time of start-up along the open circuit described above is due to the fact that the pressure in the tank 1 is above atmospheric. Due to the filling of the lines of the cooling system circuit with the coolant from the tank 1, the primary medium that was there before the valve 2 was opened is displaced. Next, at a certain point in time, the valves 2, 9 are closed. The common shaft 7 with the turbine 3 and pump 5 located on it continues to rotate by inertia (like a flywheel). The pump 5 due to the stored mechanical energy continues to pump the coolant in the lines located between the pump 5 and the check valve 8, and at the same time continues to suck the coolant from the lines in the section containing the turbine 3 and the heat exchanger 4. At the same time, the pressure in the section containing the cooling path supersonic part of the nozzle, grows, and in the area containing the turbine 3 and the heat exchanger 4, falls. The indicated trends in pressure changes are amplified by a decrease in the temperature of the coolant in the heat exchanger and an increase in its temperature in the heated cooling path of the supersonic part of the nozzle. As a result, a predetermined pressure drop occurs on the check valve 8, it opens, a closed coolant circulation loop appears, after which the cooling system enters a stationary mode. In the stationary mode of operation, the coolant cools the wall of the supersonic part of the nozzle, turning into steam, then the resulting steam "works" on the turbine 3 and enters the heat exchanger 4, where, transmitting thermal energy to the refrigerant, it liquefies and enters the pump 5, is compressed in it and again fed into the cooling path of the supersonic part of the nozzle 6. One of the fuel components is used as a refrigerant in the heat exchanger. Thus, in the stationary mode of operation, the heat supplied to the coolant in the cooling path is partially converted into mechanical energy on the turbine by the rotary pump, and the heat not used on the turbine is transferred to the coolant in the heat exchanger, which, as a fuel component, then enters the chamber and burns.

In the device for regenerative cooling of the supersonic part of the nozzle (Figure 2) there is no exhaust pipe with a valve - instead of it a receiver 10 is installed. In stationary mode, the proposed cooling device works in exactly the same way as discussed above (Figure 1). Differences are in the startup procedure. The cooling device from the very beginning is a closed loop filled with low pressure coolant vapors. After the valve 2 is opened, the movement of the coolant along the circuit does not occur due to bleeding the coolant from the coolant tank 1 to the atmosphere, but due to filling the tank of the receiver 10. The rest of the startup process does not differ from the above.

When a cryogenic fuel component is used as a coolant, a substance with a solidification temperature lower than the temperature of the coolant in the heat exchanger can be used as a coolant, which will exclude the possibility of freezing of the coolant, for example, coolant - liquid oxygen, coolant - methane. An advantage of the proposed device for regenerative cooling of the supersonic part of the nozzle is that it becomes possible to choose a more efficient cooler for the nozzle that is not part of the fuel. Currently, as part of the Federal Space Program, the study is underway of engines made according to the scheme with afterburning of reducing gas. Also in Russia, the recently created operating oxygen-methane engine demonstrator C5.86 with afterburning of reducing gas has been created. The proposed device can significantly improve the reliability of engines of this type.

Information sources

1. RF patent No. 2166661 "Method of operation of a liquid propellant rocket engine with a turbopump supply of oxygen-methane fuel", publ. 05/10/2001.

2. RF patent No. 2209993 "Method of operation of a liquid propellant rocket engine with a turbopump supply of oxygen-methane fuel", publ. 08/10/2003.

3. RF patent No. 2205288, "The cooling system of the combustion chamber of a liquid rocket engine", publ. 05/27/2003.

4. DF Slesarev, VI Tararyshkin and others. "Features of the organization of cooling nozzles of marching rocket engines, made according to the scheme with afterburning of regenerative generator gas", publ. Flight magazine, No. 7,2011.

Claims (1)

A device for regenerative cooling of the supersonic part of a nozzle of a liquid propellant rocket engine, including a heat carrier tank with a valve and a coolant circulation circuit, consisting of a cooling path for the supersonic part of the nozzle, turbine, heat exchanger, pump and the mains connecting them, characterized in that a check valve is installed in the coolant circulation circuit on the line connecting the output of the cooling path of the supersonic part of the nozzle to the turbine inlet, and the coolant tank with the valve is connected to this line in the section between the check valve and the turbine, in addition, in the section of the line between the exit of the cooling path the supersonic part of the nozzle and the non-return valve is connected to an exhaust pipe with a valve or receiver.

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