RU2563641C2 - Hybrid aerospace rocket ramjet - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: claimed engine comprises rocket engine running on fuel composed of aluminium nanopowder sized to not over 25 nm in liquid aqueous phase and ramjet combined therewith and running on molecular hydrogen formed at combustion of aluminium nanopowder. Claimed engine comprises cylindrical shell ring, central axially symmetric or wedge-like body fitted in said shell ring, combustion chamber to combust aluminium nanopowder in water vapours. One end of said cylindrical shell ring has inlet for atmospheric airflow and discharge nozzle at opposite end. System of shock waves is developed at edges of said central axially symmetric or wedge-like body. Said combustion chamber to combust aluminium nanopowder in water vapours doubles as the chemical reactor for production of hydrogen and is located in wedge-like body and coupled therewith at outlet and with nozzle suitable for formation of combustion zone at interaction of hydrogen efflux with incoming airflow. Molecular hydrogen and oxidizer in combustion zone exist in stoichiometric relationship (fuel-to-oxidizer ratio ϕ=1), or form lean mix (with ratio ϕ<1).

EFFECT: higher efficiency of fuel chemical energy conversion into thermal power, expanded range of flying conditions.

3 cl, 1 dwg

Description

The invention relates to the field of engines for the aerospace industry.

A well-known aerospace rocket engine designed for space planes acting as shuttles between the Earth and space, consisting of thermonuclear and electronic electric jet engines (ERE) (RF patent No. 2140014). During start-up, a thermonuclear electric propulsion and electronic electron propulsion are operating. A thermonuclear propulsion engine is used to launch acceleration, and an electronic propulsion engine plays the role of a ramjet to the upper atmosphere. Thermonuclear electric propulsion engines start to work there and the space plane enters outer space.

Known ramjet engine with a mechanical system for supplying fluidized powdered fuel (patent for utility model of the Russian Federation No. 10222 from 08/17/1998, IPC ⁶ F02K 9/70), consisting of a system for supplying powdered fuel, a jet turbine and a gas-permeable piston. The device has a bulky design, including a turbine with a screw and a single combustion chamber, which does not provide the required completeness of fuel combustion.

Known combined jet engine (RF patent No. 2280778). The jet thrust of a turbojet engine for launching, accelerating and flying the aircraft is created by the expiration of compressed gas through the jet nozzle by compressing the air with a compressor driven by a gas turbine, burning fuel in the combustion chamber and further burning fuel in the combustion chamber located in front of the jet nozzle . After the aircraft is accelerated to the required flight speed, a ramjet engine is used, previously compressing the air in it due to the high-speed pressure and feeding it into the exhaust combustion chamber from its sides, bypassing the compressor, creating radially directed flows. An additional compression of the flow is performed in the output combustion chamber, for which purpose flows of incoming compressed from the periphery to its center air flow due to the high-pressure head are created, pushing them in the central part of the output combustion chamber with mutual braking and converting their kinetic energy into additional compression. Additional fuel combustion is carried out in the area of increased air compression.

Known ramjet engine on a metal powdered fuel containing a start system, a fuel supply system comprising a fuel tank with a metal powdered fuel and a perforated piston and a combustion chamber. A gas generator is installed in the engine casing, which enables the start-up of the powder metal fuel supply system, while the combustion chamber consists of a prechamber with a length of at least 300 mm, with a dispenser located at the inlet of the chamber supplying fuel with a predetermined flow rate, an igniter and a final fuel afterburner, and the final afterburner have air supply channels (RF patent No. 2439358).

Known technical solutions do not fully realize the energy resources of the fuel, do not allow to obtain the maximum efficiency of the combustion products, and therefore, do not allow to solve the problem of creating an economical engine for aerospace aircraft operating in a wide range of flight modes.

In the aerospace industry, an important direction in the creation of new aircraft is to reduce fuel consumption and increase engine specific thrust. To increase the specific characteristics of jet engines, it is necessary to use fuels with either a higher calorific value or with a higher efficiency of combustion products.

This is due to the fact that with an increase in the flight speed at the same altitude, the drag of the aircraft increases approximately in proportion to the square of the flight speed, while the specific air flow through the engine path is proportional to the speed; therefore, jet engines designed for high supersonic speeds flight, should have a higher specific thrust compared with engines operating at moderate speeds. If the required specific thrust is sufficiently large, then the greatest profitability is achieved if the fuel is used with the maximum efficiency of the combustion products. Therefore, it is necessary to use a higher-energy fuel and the choice of a constructive solution that will maximize the energy resources of fuel.

The basis of the invention is the creation of a new economical engine for aerospace aircraft operating in a wide range of flight modes.

The technical result is an increase in the efficiency of converting the chemical energy of fuel into thermal energy and increasing the efficiency of combustion products.

Another technical result is the expansion of the range of flight modes (for the flight Mach number from 0 to 24).

The problem is solved by the fact that a hybrid rocket-ramjet air-propelled aerospace engine is proposed, including a rocket engine (RD) on composite fuel consisting of aluminum nanoparticles no larger than 25 nm and liquid water, operating both at launch and during flight the aircraft, and the ramjet engine combined with it, ramjet engine operating on molecular hydrogen generated during the combustion of aluminum nanopowder in water vapor, which is used together with the taxiway as at the stage the acceleration of the apparatus, and in cruising mode, characterized in that it includes an external shell, forming an air intake and an air duct, a central axisymmetric or wedge-shaped body located in the shell, a supersonic exit nozzle, a combustion chamber for burning aluminum nanopowder in water vapor, which is also a chemical a hydrogen producing reactor located in the central body and conjugated to a supersonic nozzle of the central body, at the exit of which a combustion zone is formed, which is formed during the interaction of hydrogen flowing out of the nozzle and the air flow coming from the duct.

It is extremely advisable if in the combustion zone molecular hydrogen and an oxidizing agent are in a stoichiometric ratio (fuel / oxidizing ratio ϕ = 1), or they form a poor mixture (ϕ <1).

A hybrid ramjet ramjet aerospace engine is proposed that combines a conventional rocket engine (RD) and ramjet (ramjet). Moreover, the latter uses molecular hydrogen as a fuel on a cruise flight mode, and atmospheric air as an oxidizer. These are the products of the oxidation of a mixture of aluminum with water vapor, which runs a rocket engine.

The figure shows a schematic diagram of an engine according to the invention.

The outer part of the engine, according to the invention, is a shell of ramjet of circular or rectangular cross-section 1. Air 10 enters the air intake and is compressed to the required degree of compression (after the air intake pressure is P = 0.1-3 atm, the temperature depending on the speed and height of flight 500-1400 K) the central body 3 (axisymmetric or wedge-shaped), on the front edges of which at supersonic flight speed a system of shock waves 13 is realized.

Inside the central body 3, there is a combustion chamber of a rocket engine 5, which is connected to a nozzle 4, which exits into the internal path of an air-jet jet engine 1. Combustion chamber 5 is also a constant pressure chemical reactor for producing molecular hydrogen from aluminum and water vapor used in taxiways as fuel. An emulsion of non-oxidized nanoparticles of aluminum 6 is fed into the combustion chamber 5 in a liquid aqueous medium through nozzles 14 and is dispersed as a result of spraying, and then evaporates, and already water vapor, entering the oxidation reaction with nano Al in zone 8, generate primary combustion products 9 - molecular hydrogen and aluminum oxides, their temperature in the combustion zone reaches 3700 K. The most temperature-stressed object in the engine is the combustion chamber 5, therefore its walls are externally cooled by liquid water 7, which flows against the flow of combustion products to sunki 15 through film cooling technology. In addition to actually cooling the walls, this will improve the fineness of the atomization of water and, essentially, atomized superheated water vapor will be supplied to the combustion chamber 5 through nozzles 15, which will also react with non-oxidized Al nanoparticles. The combustion products 9, flowing out of the nozzle 4, create draft during expansion in the nozzle, and the temperature of the combustion products drops to 1300-1500 K. Further, downstream, the combustion products containing molecular hydrogen (molar fraction of γ _H2 = 75%) and Al ₂ O ₃ in the condensed phase are mixed with air passing through the ramjet path, either in the stoichiometric ratio (fuel / oxidizer ratio ϕ = 1), or with the ratio ϕ <1 (lean mixture), and, having formed the mixture, self-ignite and burn out combustion zone 11, forming secondary combustion products 12 (they consist mainly m of H ₂ O, Al ₂ O ₃ in the liquid phase and N ₂ ) with a temperature of 2400-2700 K. The combustion products 12 expire from the nozzle 2 and create additional thrust (to the thrust RD).

In fact, the oxidation reaction of a stoichiometric mixture of aluminum with water 2Al + 3H ₂ O => Al ₂ O ₃ + 3H ₂ proceeds with the formation of hydrogen and the release of a significant amount of heat Q = 481 kJ / mol. The heat generated is converted into the kinetic energy of the primary combustion products, and the jet stream creates traction. To ensure the effective operation of the ramjet, it is necessary that a certain amount of air flow through the engine path be set and that a sufficient amount of fuel can be burned and traction provided, i.e. such an engine that can only work starting at a certain flight speed. To accelerate the apparatus to the desired speed, a rocket engine is used, in the combustion chamber of which the oxidation reaction 2Al + 3H ₂ O => Al ₂ O ₃ + 3H ₂ takes place. As a result of the contact of non-oxidized aluminum with water vapor, the particles are oxidized, i.e. covered with an oxide film that forms very quickly and prevents further oxidation (the boiling point of the oxide film is 2380 K). For certain sizes of Al particles that have not yet been oxidized (radius less than 25 nm), the oxidation reaction of the surface of the particles occurs with such a high heat that the particle will not have time to transfer heat to the outer space and the aluminum inside the particle will begin to boil and destroy the oxide shell during strong expansion. In this case, aluminum will atomize and react with H ₂ O. In this case, in contrast to the combustion of micrometer-sized particles, aluminum is almost completely burnt in water vapor. Moreover, in the combustion products, liquid Al ₂ O ₃ particles are formed through the mechanism of homogeneous nucleation and, as calculations have shown, during the stay of the mixture in the combustion chamber 5 and nozzle 4, their size does not have time to significantly increase, and the bulk of the liquid Al ₂ O ₃ particles has size 40-50 nm. Such particles have short thermal and dynamic relaxation times (~ 10 ^-7 -10 ^-6 s) and do not lead to noticeable losses in the specific impulse due to different velocities and temperatures of the gas-phase and liquid-phase continua (loss of two-phase). At the same time, during combustion of Al particles of a micrometer size, the kinetic rather than diffusive (substantially slower) combustion mode is not realized and the particles in this case do not burn out completely (the smallest particles with a size of 5-15 nm remain). In this case, the formation of the liquid phase Al ₂ O ₃ in the combustion products occurs due to heterogeneous condensation and the resulting particles reach micron sizes (1-20 microns). Such particles have very large times of thermal and dynamic relaxation, which leads to large losses in two-phase state (it is impossible to convert all the energy released during combustion into a specific impulse). Therefore, it is proposed to store non-oxidized Al nanoparticles with a radius of less than 25 nm in a liquid-free H ₂ O phase on the fuel lines and supply through the fuel lines, and only then, when supplied through nozzles, atomize water so that the oxidation and atomization of aluminum followed by oxidation of atomic aluminum in water vapor occurred in the combustion chamber of the taxiway.

The efficiency of primary combustion products at P = 1 atm was calculated. The primary combustion products of Al with H ₂ O at the outlet of chamber 5 are a mixture of H ₂ and Al ₂ O ₃ particles in the liquid phase 3.53 / 1, because obtained by burning a poor mixture of aluminum and water vapor. The efficiency of the combustion products Ae is determined by the expression R · ΔTe / μ, where R is the gas constant, ΔTe = T _a -T _c = 3700K-1200K = 2500K is the adiabatic combustion temperature T _a minus the temperature of the combustion products T _c in the outlet section of the nozzle 4, µ = 24 g / mol is the molecular weight of the primary combustion products. Under these conditions, Ae = 870 kJ / kg, which is approximately 10% higher than the values of Ae realized by burning pure aluminum or hydrogen in air at a stoichiometric ratio of fuel / oxidizer ϕ = 1. The efficiency of kerosene combustion in air with stoichiometry, compared with the calculated case, is even lower - almost 1.3 times.

Similarly, the calculated efficiency of secondary combustion products. Secondary products are a mixture of H ₂ O / Al ₂ O ₃ (g) / N ₂ = 24.5 / 8.6 / 56.8, obtained under the condition of stoichiometric combustion of primary products in air. At ΔTe = T _a -T _c = 2630K-250K = 2380K and µ = 30 g / mol, it amounted to 660 kJ / kg, which approximately corresponds to the efficiency of kerosene combustion products in air. When working together RD and ramjet, it is necessary to take into account the performance of both primary and secondary combustion products, which in total amounts to 1530 kJ / kg, which is approximately 2.3 times more than that of kerosene combustion products in air.

Due to the use of nano Al + H ₂ O fuel in the RD combustion chamber, it is possible to increase the efficiency of the direct-flow ramjet engine, compared to a conventional ramjet, because if the required specific thrust is large enough, then the greatest economy is achieved if fuel with maximum efficiency of the combustion products and the specific impulse of the proposed RPD using the combustion of Al nanoparticles (with a radius of R <25 nm) in water vapor is several times greater than the specific impulse of ramjet engine operating on kerosene, although the calorific value of Al is It is slightly lower than that of pure kerosene. Despite the fact that the proposed RPD at high specific thrusts actually uses a fuel mixture enriched with fuel, its specific thrust is nevertheless higher than the specific thrust of a traditional liquid propellant liquid propellant rocket engine operating on a kerosene / air mixture.

External ramjet 1 is effective only at high flight speeds (M> 0.8). Therefore, the start is carried out using a taxiway, which, of course, is less economical than a conventional gas turbine engine, since for a taxiway not only fuel, but also an oxidizing agent must be taken with you. In any case, the initial phase of the flight is usually short-lived, and at the stage of acceleration and the main cruising mode it is already included in the work and creates additional thrust ramjet thrust for which fuel is generated in the RD combustion chamber and the oxidizer is taken from the atmosphere. An important advantage of the proposed hybrid RPD is the ability to use additional traction from a rocket engine in a cruise flight mode. In addition, the layout with the internal placement of the booster taxiway is more compact and reduces drag as compared to the conventional separate layout (sometimes even to reduce drag and not carry ballast, the booster engine is shot off, which is clearly not suitable for reusable aerospace systems).

The invention can be used both for guided missiles designed to fly at high supersonic speeds at moderate altitudes (~ up to 50 km), and for aerospace systems, with the possibility of flying both in the atmosphere and in near space.

Claims (3)

1. A hybrid ramjet ramjet aerospace aerospace engine, characterized in that it includes a rocket engine (RD) in the form of aluminum nanopowder with a size of not more than 25 nm in the liquid aqueous phase, designed to create traction at launch and at the initial stage of flight, and for the production of fuel consumed by the combined-flow ramjet engine with a molecular hydrogen cruise generated during the combustion of aluminum nanopowder, and characterized in that it contains a cylindrical a shell, at one end of which there is an inlet for atmospheric air flow, and at the other — an outlet nozzle, a central axisymmetric or wedge-shaped body, on the edges of which a system of shock waves is implemented, located in the shell, a combustion chamber for burning aluminum nanopowder in water vapor, which is simultaneously a chemical reactor for producing hydrogen, placed in a wedge-shaped body, coupled to it at the outlet and with a nozzle adapted to form a combustion zone during the interaction of hydrogen flowing from the nozzle and flowing air stream. 2. The hybrid ramjet ramjet aerospace engine according to claim 1, characterized in that in the combustion zone molecular hydrogen and an oxidizing agent are in a stoichiometric ratio (fuel / oxidizer ratio ϕ = 1), or they form a lean mixture (ratio ϕ <1 ) 3. A hybrid ramjet ramjet aerospace engine according to claim 1 or 2, characterized in that it operates at a speed equal to the Mach number M = 0-24.

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