RU2714582C1 - Method for arrangement of working process in straight-flow air-jet engine with continuous-detonation combustion chamber and device for implementation thereof - Google Patents

Abstract

FIELD: physics.

SUBSTANCE: invention relates to methods of operating a working process in air-jet engines with continuous-detonation combustion and devices for their implementation, particularly for high-speed unmanned aerial vehicles. Method of working process organization in straight-flow air-jet engine with continuous-detonation combustion includes acceleration of aircraft to supersonic speed providing for beginning of autonomous flight of aircraft with such engine, partial deceleration of incoming supersonic air flow in skew races of seal and in wall boundary layer before arrival in annular combustion chamber, continuous supply of fuel to mixing zone with air, formation of detonation-capable mixture of fuel and air. Further, continuous-detonation combustion of fuel mixture is initiated, flow of detonation products from annular combustion chamber through annular nozzle with supersonic speed with formation of jet stream and creation of reactive thrust. Approaching supersonic air flow is first partially inhibited in skew jumps of the seal and in the wall boundary layer, and then accelerated in the fan of rarefaction waves with partial recovery of parameters of incoming supersonic air flow and enters the straight-flow air-jet engine in the form of weakly retarded supersonic air flow. One part of air is directed into annular combustion chamber. Other part, including wall boundary layer, is directed around annular combustion chamber to provide cooling of walls of annular combustion chamber and to prevent gas-dynamic effect of continuous-detonation combustion of fuel mixture and air in the annular combustion chamber for the course of weakly retarded supersonic air flow at the inlet of the straight-flow air-jet engine. Method is implemented in the device including a supersonic air intake, a central body, an annular combustion chamber with a belt of nozzles of fuel supply, a gas-dynamic insulator located between the annular combustion chamber and the external wall of the rear cone of the central body.

EFFECT: invention provides the possibility of autonomous flight with low Mach number.

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Description

Technical field

The invention relates to methods for organizing a working process in ramjet engines with continuous detonation combustion and devices for their implementation, intended, in particular, for high-speed unmanned aerial vehicles.

State of the art

Modern high-speed unmanned aerial vehicles, as a rule, are equipped with small-sized turbojet or ramjet engines. It is known that the efficiency of small turbojet engines during flight at a speed exceeding the speed of sound decreases sharply (VV Rostopchin. Micro-turbojet engine for unmanned aerial vehicles, Moscow, Central Research Institute of ARKS, 2004). Thus, one of the most important characteristics of the efficiency of turbojet engines — the specific impulse equal to the ratio of thrust to the second weighted fuel consumption — is approximately 1300 s with a Mach number of flight 1 and decreases to ~ 800 s with a Mach number of 1.25.

Existing ramjet engines operating on a thermodynamic cycle with continuous deflagration fuel combustion at constant pressure are effective in the range of Mach flight numbers of an aircraft from about 2 to 6. One of the most important characteristics of an aircraft with a ramjet engine is the minimum number Mach, in which an autonomous flight of an aircraft begins after a launch acceleration provided by various auxiliary accelerating devices with water. As indicated above, for existing aircraft with a ramjet engine, this Mach number is approximately 2. At a lower flight speed, the incoming air flow has a low braking pressure, which narrows the area of stable operation of the air intake device and reduces the thrust characteristics of the engine to critical values, which autonomous flight of the aircraft becomes impossible. To accelerate the aircraft to this speed, auxiliary acceleration devices, such as rocket engines, are usually used. The rocket engines used for launch acceleration of the aircraft have relatively low specific thrust characteristics, which leads to an increase in the starting mass and launch dimensions of the aircraft. So, for example, an assessment according to the well-known formula of Tsiolkovsky K.E. shows that to accelerate the aircraft to Mach number M = 1, an auxiliary acceleration device in the form of a rocket engine is required, the fuel mass of which is about 10-20% of the mass of the entire aircraft. To accelerate the aircraft to the Mach number M = 2, a fuel mass of about 20-40% of the mass of the entire aircraft is required. It is seen that in order to reduce the starting mass of the aircraft, it is necessary to reduce the Mach number of the beginning of the operation of the ramjet engine.

The traction characteristics of ramjet engines have actually reached limit values and their further increase is problematic. Therefore, to obtain a significant increase in the efficiency of ramjet engines, the possibility of using a thermodynamic cycle with continuous detonation combustion of fuel in an annular combustion chamber is considered (Frolov S.M., Zvegintsev V.I., Ivanov V.S., Aksenov V.S., Shamshin I.O., Vnuchkov D.A., Nalivaichenko D.G., Berlin A.A., Fomin V.M. Continuous detonation combustion of hydrogen: results of tests in a wind tunnel.Physics of Combustion and Explosion, 2018, v. 54 No. 3, pp. 116-123). It is known that such a cycle has a higher efficiency (Frolov S.M., Barykin A.E., Borisov A.A. Thermodynamic cycle with detonation fuel combustion. Chemical Physics, 2004, Volume 23, No. 3, p. 17 -25). In addition, in comparison with continuous deflagration combustion, continuous detonation fuel combustion has several advantages. Thus, with continuous deflagration combustion, energy is released over the entire cross section of the engine’s combustion chamber and creates a backpressure against the incoming air flow, which makes it difficult to organize the working process at a relatively low flight speed of the aircraft and leads to disruption of the intake device. During continuous detonation combustion, energy release occurs in a narrow reaction zone in one or more detonation waves that continuously circulate in the annular combustion chamber. Since the detonation wave at each moment of time occupies only a small part of the cross section of the combustion chamber, this allows you to organize the working process at a relatively low flight speed of the aircraft without disrupting the operation of the air intake device. Another important advantage of continuous detonation combustion is the possibility of organizing a workflow with a high degree of combustion in an engine with a relatively small longitudinal size corresponding to the length of a self-sustaining detonation wave. This can significantly reduce the loss of total pressure in the engine path compared to a continuous deflagration combustion engine. The disadvantages of continuous detonation combustion include its unsteady nature, however, due to the high speed of detonation waves continuously circulating in the annular combustion chamber, the pulsations of the flow parameters in the jet stream of the engine are high-frequency (several kilohertz), and the working process is quasi-stationary.

In our works (Dubrovsky A.V., Ivanov V.S., Zangiev A.E., Frolov S.M. Three-dimensional numerical modeling of the characteristics of a direct-flow air-reactive power plant with a continuous detonation combustion chamber under conditions of supersonic flight. Chemical Physics, 2016, Vol. 35, No. 6, pp. 49-63; Frolov S.M., Zvegintsev VI, Ivanov VS, Aksenov VS, Shamshin IO, Vnuchkov DA, Nalivaichenko DG, Berlin AA, Fomin VM, Shiplyuk AN, Yakovlev NN Hydrogen -fueled detonation ramjet: Wind tunnel tests at approach air stream Mach number and stagnation temperature 1500 K. International Journal of Hydrogen Energy, Vol. 43, pp. 7515-7524; Frolov S.M., Zvegintsev V.I., Ivanov BC , Aksenov BC, Shamshin I.O., Granddaughters D.A., Nalivaichenko D.G., Berlin A.A., Fomin V.M. Continuous detonation combustion of hydrogen: results of tests in a wind tunnel.Physics of Combustion and Explosion, 2018, v. 54, No. 3, p. 116-123) describes the results of calculations and fire tests of a hydrogen continuous detonation ramjet engine with an annular combustion chamber in an incident supersonic air flow. A continuous detonation workflow with positive effective thrust was experimentally implemented in the range of Mach numbers of the incoming air flow from 4 to 8 at a braking temperature of 300 K to 1500 K, and high values of full thrust (up to 2200 N) were recorded in the fire tests of the developed engine demonstrator and specific impulse (up to 3600 s).

In view of the distinguishing features of detonation combustion noted above, it can be expected that the minimum Mach number at which an autonomous flight of an aircraft equipped with a ramjet engine will start will be less for a continuously detonating engine than for an engine running on continuous deflagration fuel combustion .

The known method and device proposed in the patent of the Russian Federation 2585328 C2, F02K 7/02 (2006.01), F02K 7/14 (2006.01), 05/27/2016. The detonation-deflagration pulsating ramjet engine contains a supersonic air intake, a continuous fuel supply system, a lattice plate damper of detonation waves, located so that it receives a well-mixed combustible mixture, a combustion chamber and an exhaust nozzle. The supersonic air intake inhibits the oncoming high-speed supersonic air flow to Mach numbers M = 3-4. The grating plate damper comprises one or more plates located along the axis of the engine flow path. The transverse size of each channel formed by the absorber plates is smaller than the transverse size of the cells generated during the combustion of a detonation wave, moving against the flow and incident on the same damper, which stops and dampens the propagation of the detonation wave when it enters the narrow channels of the damper, and shock waves, arising from the extinction of the detonation wave, it carries out from the channels into the combustion chamber by a supersonic flow, preventing them from breaking the flow of the incoming flow and restricting the movement of detonation and shock waves n absorber part and the combustion chamber, providing the transition from combustion to detonation, whereby arranged continuous unsteady combustion in a dynamically pulsating (emerging and fading) detonation wave fronts and slow combustion. The technical result is an increase in thrust and an extension of the range of flight speeds to Mach numbers M = 5-8 with a decrease in the heat stress of the engine path compared to a ramjet engine and a ramjet engine with supersonic combustion. The method and device have disadvantages. Firstly, the elements of the grating plate damper of detonation waves clutter the engine flow path and lead to large losses of total pressure. Secondly, the estimated minimum Mach number of autonomous flight exceeds M = 3-4 due to the inhibition of air flow in a supersonic air intake, which requires the use of powerful auxiliary booster devices that can provide such a high flight speed.

The known method and device proposed in the patent of the Russian Federation 2573427 C2, F02K 7/02 (2006.01), F02K 7/14 (2006.01), 05/27/2016. The method of burning the air-fuel mixture to create jet thrust in a ramjet engine with a spin detonation wave is that the incoming high-speed flow is slowed down to Mach numbers in the range from M = 3 to M = 4 in a supersonic two-stage air intake with a dull central body. Next, fuel is fed into the stream, the resulting air-fuel stream is twisted into the well-mixed combustible mixture, braked to the subsonic axial velocity component, ignition of the swirling well-mixed fuel-air mixture is initiated, and burnt in a spin detonation wave. Detonation and shock waves propagating against the flow are quenched by the supersonic flow of the air-fuel mixture. The combustion products formed during combustion are directed to create jet thrust of the engine. A ramjet engine with a spin detonation wave for high-speed flights contains a supersonic two-stage air intake with a dull central body, a drainage system for the entropy and boundary layers, fuel pylons with nozzles for supplying fuel to the incoming air flow, the crowns of which are made and arranged so that they continue braking and spin the resulting air-fuel flow, annular lattice damper of detonation and shock waves, axisymmetric annular nozzle, I have its flared outer shroud and a central body with a base cut. The annular lattice damper of detonation and shock waves contains annular lattice partitions forming channels for braking and turning the air-fuel flow to the subsonic axial velocity component while maintaining the supersonic velocity in the absorber channels. At the outlet of the damper there is an annular detonation combustion chamber, the initial inner radius of which is less than the inner radius of the damper rings. An annular lattice is located at the outlet of the combustion chamber, which straightens the exit stream. The invention is aimed at intensifying the rate of chemical reactions of combustion and energy release due to spin detonation combustion of a well-mixed air-fuel mixture. The method and device have disadvantages. Firstly, the elements of the annular lattice absorber of detonation and shock waves clutter the engine flow path and lead to large losses of total pressure. Secondly, the estimated minimum Mach number of autonomous flight exceeds M = 3-4 due to air flow inhibition in a supersonic two-stage air intake with a dull central body, which requires the use of powerful auxiliary accelerating devices capable of providing such a high flight speed.

Closest to the proposed invention in technical essence are the method and device proposed in the article Frolova S.M., Zvegintseva V.I., Ivanova BC, Aksenova BC, Shamshina I.O., Vnuchkova D.A., Nalyvaychenko D.G. ., Berlin A.A., Fomina V.M. Continuous-detonation combustion of hydrogen: wind tunnel test results. Combustion and Explosion Physics, 2018, v. 54, No. 3, p. 116-123. The prototype method includes accelerating an aircraft with a ramjet engine installed on it for continuous detonation combustion to a supersonic flight speed with a Mach number of at least M = 4, ensuring the start of an autonomous flight of the aircraft, partial braking of the incident supersonic air flow in oblique jumps seals in the parietal boundary layer before entering the annular combustion chamber, continuous supply of fuel to the mixing zone with air, formation of detonation special mixture of fuel and air, the initiation of continuous-detonation combustion of the fuel mixture, the expiration of detonation products from the annular combustion chamber at a supersonic speed through the annular nozzle with the formation of a jet stream. The prototype device contains an axisymmetric supersonic three-hop air intake designed for partial braking of the incident supersonic air flow in three oblique shock waves, an axisymmetric central body, an expanding axisymmetric annular combustion chamber with a belt of radial fuel injectors and an expanding annular nozzle. The central body is formed by front and rear cones connected by bases. The prototype method and the prototype device have disadvantages:

(1) the estimated minimum Mach number of autonomous flight exceeds M = 4, which requires the use of powerful auxiliary booster devices capable of providing such a high flight speed;

(2) a limited amount for the placement of functional blocks (payload, control system elements, fuel supply, etc.);

(3) To protect functional blocks from high temperatures and vibrations caused by continuous detonation combustion, special measures and approaches are required.

Disclosure of Invention

The objective of the invention is the creation of such a method of organizing a workflow in a ramjet engine with continuous detonation combustion, which will ensure the start of an autonomous flight of an aircraft with such an engine with a low Mach number, provide sufficient volume to accommodate functional blocks and protect functional blocks from high temperatures and vibrations caused by continuous detonation combustion using standard measures and approaches.

The objective of the invention is the creation of a device for implementing a method of organizing a work process in a ramjet engine with continuous detonation combustion, which will ensure the start of an autonomous flight of an aircraft with such an engine with a low Mach number, will provide sufficient volume to accommodate functional blocks and protect functional blocks from high temperatures and vibrations caused by continuous detonation combustion, using standard measures and approaches.

The solution to this problem is achieved by the proposed:

- a method of organizing a working process in a ramjet engine with continuous detonation combustion, including accelerating the aircraft to supersonic speed, ensuring the start of an autonomous flight of an aircraft with such an engine, partial braking of the incident supersonic air flow in oblique shock waves and in the wall boundary layer before entering the annular combustion chamber, a continuous supply of fuel to the mixing zone with air, the formation of detonation mixture of fuel and air, initiation of continuous detonation combustion of the fuel mixture, outflow of detonation products from the annular combustion chamber through the annular nozzle at a supersonic speed with the formation of a jet stream and the creation of reactive thrust in which the incident supersonic air flow is first partially inhibited in oblique shock waves and in the near-boundary boundary layer, and then accelerates in a fan of rarefaction waves with a partial restoration of the parameters of the incident supersonic air flow and from enters a ramjet in the form of a slightly inhibited supersonic air flow, with one part of the air being directed into the annular combustion chamber to form a detonation-friendly mixture of fuel and air at such a distance from the inlet section of the annular combustion chamber so that continuous detonation combustion in the annular chamber combustion did not lead to the displacement of the mixture of fuel and air, as well as hot detonation products from the annular combustion chamber through its inlet section, and the other part, including The flashing near-wall boundary layer is directed to bypass the annular combustion chamber to provide cooling of the walls of the annular combustion chamber and to prevent the gas-dynamic effect of continuous detonation combustion of the fuel-air mixture in the annular combustion chamber on the flow of weakly inhibited supersonic air flow at the inlet of the ramjet engine .

- a device comprising a supersonic air intake, a central body with front and rear cones, an annular combustion chamber with a belt of fuel nozzles and an annular nozzle in which the central body contains an additional section between the front and rear cones, and a supersonic air intake communicates with the annular combustion chamber and with gas-dynamic insulator, made in the form of a bypass channel and located between the annular combustion chamber and the outer wall of the rear cone of the Central body, and the front chrome a wall that separates the annular combustion chamber of a gas-dynamic insulator shifted deep into the supersonic air intake, and the rear edge of the wall separating the combustion chamber from the annular gas-dynamic insulator shifted deep into the annular nozzle.

An additional section of the central body may have a conical, cylindrical, or other shape, providing a continuous flow of supersonic air flow.

The supersonic air intake can be made according to the internal compression scheme due to the profiling of its internal path or according to the compression scheme in the attached oblique shock waves.

The belt of fuel injectors in the annular combustion chamber is located at such a distance from the inlet section of the annular combustion chamber so that continuous detonation combustion in the annular combustion chamber does not displace the mixture of fuel and air, as well as hot detonation products from the annular combustion chamber through its inlet section .

The rear cone of the axisymmetric central body can be truncated and have the form of a straight cone or cone with a profiled side surface.

The annular nozzle is preferably made in the form of an expanding annular channel in order to ensure the full expansion of the detonation products coming from the annular combustion chamber and the air flow coming from the gas-dynamic insulator to atmospheric pressure without shock waves.

Brief Description of the Drawings

In FIG. 1 shows a diagram of the inventive device. In FIG. 1 is designated: 1 - axisymmetric central body, 2 - axisymmetric supersonic air intake, 3 - axisymmetric annular combustion chamber, 4 - gas-dynamic insulator, 5 - annular nozzle, I - front cone of the central body, II - rear cone of the central body, III - additional section central body, LT - streamline, KSU - oblique shock wave, BP - fan of rarefaction waves, C - wall, PKS - front wall edge, ZKS - trailing edge of the wall, PF - belt for fuel injectors, T - fuel.

In FIG. 2 shows the calculated dependence of the specific impulse (fuel) of the proposed device in the range of Mach numbers of the incoming air flow from 1.7 to 2.2.

The implementation of the invention

The device consists of a supersonic air intake (2), a central body (1), an annular combustion chamber (3) with a belt of fuel injectors (PF) and an annular nozzle (5). The central body (1) has an additional section (III) located between the front (I) and rear cones (II) and made in the form of a truncated cone, with its smaller base facing the rear cone (II) of the central body (1). The supersonic air intake (2) communicates with the annular combustion chamber (3) and with the gas-dynamic insulator (4), made in the form of a bypass channel and located between the annular combustion chamber (3) and the outer wall of the rear cone (II) of the central body (1), and the leading edge (PKS) of the wall (C) separating the annular combustion chamber (3) from the gas-dynamic insulator (4) is displaced deep into the supersonic air intake (2), and the trailing edge (ZKS) of the wall (C) separating the annular combustion chamber (3) from the gas-dynamic insulator (4), is displaced deep annular nozzle (5). The device also includes functional blocks: a detonation initiation system, a control system and a fuel supply system (not shown in FIG. 1).

The proposed device operates as follows.

The device accelerates in any known manner to supersonic flight speeds with a minimum Mach number, which provides the start of an autonomous flight of an aircraft. On the front cone (I) of the central body (1) there is an oblique shock wave (KSU), not included in the supersonic air intake (2), as well as a fan of rarefaction waves (BP). The incident supersonic air flow (see LT flow lines) is first partially decelerated in the oblique shock wave (CCU) and in the near-boundary boundary layer, and then accelerated in the fan of rarefaction waves (BP) with a partial restoration of the parameters of the incident supersonic air flow and enters the supersonic air intake (2) apparatus. On the front cone (I) of the central body (1) there is an oblique shock wave (KSU) that does not fall into the supersonic air intake (2), as well as a fan of rarefaction waves (BP). The incident supersonic air flow (see streamlines (LT)) is first partially decelerated in the oblique shock wave (CCU) and in the near-boundary boundary layer, and then accelerated in the fan of rarefaction waves (BP) with a partial restoration of the parameters of the incident supersonic air flow and enters supersonic air intake (2) of the jet engine in the form of weakly inhibited supersonic air flow. In a supersonic air intake (2), a weakly inhibited supersonic air stream is divided into two streams: a stream into an annular combustion chamber (3) and a stream into a gas-dynamic insulator (4). The air entering the annular combustion chamber (3) is mixed with fuel (T), which is continuously supplied to the annular combustion chamber (3) through the belt of the fuel supply nozzles (PF) using a fuel supply system. As a result, a detonative mixture of fuel and air is formed in the annular combustion chamber (3). Further, in the annular combustion chamber (3), a single initiation of continuous detonation combustion occurs in accordance with the principle described in patent WO 2014/129920 A1, Device for burning fuel in a continuous detonation wave, F23R 7/00 (2006.01), published on 08.28.2014 (authors Frolov S.M., Frolov F.S.). Continuous detonation combustion of a mixture of fuel and air in an annular combustion chamber (3) provides acceleration of combustion products towards the output section of an annular combustion chamber (3) and an annular nozzle (5) with the formation of a high-frequency quasistationary jet and thrust. It should be noted that continuous detonation combustion in an annular combustion chamber (3) is organized so that neither a mixture of fuel and air nor hot detonation products penetrate into a supersonic air intake (2). This is achieved due to the formation of a detonation-friendly mixture of fuel and air in the middle part of the annular combustion chamber (3). The air entering the gas-dynamic insulator (4) includes a wall boundary layer formed by supersonic air flow around the central body (1). This eliminates the negative effect of the boundary layer on the filling of the annular combustion chamber (3) with air. The air flow directed into the gas-dynamic insulator (4) bypassing the annular combustion chamber (3), on the one hand, provides cooling of the walls of the annular combustion chamber (3) and the rear cone (II), and, on the other hand, prevents the gas-dynamic effect of continuously detonating burning a mixture of fuel and air in an annular combustion chamber (3) during the flow of a slightly inhibited supersonic air flow at the inlet to a supersonic air intake (2). The fact is that continuous detonation combustion of a mixture of fuel and air in an annular combustion chamber (3) is accompanied by the generation of gas-dynamic disturbances in the form of regular shock waves traveling in the direction of the supersonic air intake (2), which cause an increase in pressure in the inlet section of the supersonic air intake ( 2). This can lead to disruption of the supersonic air intake (2) and to the disruption of continuous detonation combustion in the annular combustion chamber (3). Due to the fact that the leading edge (PKS) of the wall (C) separating the annular combustion chamber (3) from the gas-dynamic insulator (4) is displaced deep into the supersonic air intake (2), such regular shock waves are effective when exiting the annular combustion chamber (3) weaken and transform into weak acoustic disturbances that do not interfere with the operation of the supersonic air intake (2). It should also be noted that continuous detonation combustion in an annular combustion chamber (3) is accompanied by the generation of gas-dynamic disturbances in the form of regular shock waves traveling in the direction of the annular nozzle (5) and causing high-frequency pulsations of the jet stream and the jet propulsion generated by the engine. In addition, when the flow of detonation products from the annular combustion chamber (3) and the air stream from the gas-dynamic insulator (4) merge in the annular nozzle (5), transverse gas-dynamic perturbations and compression shocks may occur due to the pressure difference in these flows. The influence of these factors on the jet thrust generated by the engine is weakened due to the displacement of the trailing edge (ZKS) of the wall (C), which separates the annular combustion chamber (3) from the gas-dynamic insulator (4), deep into the annular nozzle (5). The profiling of the annular nozzle (5) provides a complete expansion of the detonation products coming from the annular combustion chamber (3) and the air flow coming from the gas-dynamic insulator (4) to atmospheric pressure without shock waves.

We give an example of a multidimensional gasdynamic calculation of a ramjet with continuous detonation combustion of a mixture of hydrogen and air in accordance with this invention. The calculation models used were verified on experimental data on the continuous detonation combustion of a hydrogen-air mixture in annular combustion chambers with an external diameter of 306 mm (Frolov S.M., Dubrovsky A.V., Ivanov BC // Chemical Physics. 2013. V. 32. No. 2.P. 56) and 406 mm (Frolov SM, Aksenov VS, Ivanov VS, Shamshin IO // Intern. J. Hydrogen Energy. 2015. V. 40. P. 1616.) of a continuous-detonation ramjet engine for flight conditions with a Mach number of 5 (Frolov SM, Zvegintsev VI, Ivanov VS, Aksenov VS, Shamshin IO, Vnuchkov DA, Nalivaichenko DG, Berlin AA, Fomin VM, Shiplyuk AN, Yakovlev NN Hydrogen-fueled detonation ramjet model: Wind tunnel test s at approach air stream Mach number 5.7 and stagnation temperature 1500 K, International Journal of Hydrogen Energy, 2018, Vol. 43, pp. 7515-7524) and showed good agreement with the measurements. Calculation of engine traction characteristics (traction, specific impulse, specific fuel consumption, effective traction) was carried out according to the methods described in (Dubrovsky A.V., Ivanov VS, Zangiev A.E., Frolov S.M. Three-dimensional numerical simulation of direct-flow characteristics an air-reactive power plant with a continuously detonating combustion chamber under supersonic flight conditions. Chemical Physics, 2016, Volume 35, No. 6, pp. 49-63).

The design scheme of the engine corresponded to the scheme of the inventive device shown in FIG. 1. The calculations were performed for the flow around the (external and internal) engine at sea level with an unlimited free flow of a pre-prepared stoichiometric in hydrogen-air mixture with a Mach number from M = 1.7 to M = 2.2.

The computational domain was a cylinder, the side surface of which was removed from the power plant at a sufficiently large distance to exclude its effect on the flow inside and in the vicinity of the engine. At the entrance to the computational domain, the Mach number of the incident supersonic flow and the static temperature, turbulent kinetic energy, dissipation of turbulent kinetic energy, and average mass fractions of gas components were specified. At the lateral boundaries and in the outlet section of the computational domain, atmospheric static pressure was set. All other variables (velocity, temperature, turbulent kinetic energy and its dissipation, as well as mass concentrations of the components) were extrapolated to these boundaries from the calculation domain. The static pressure and temperature in the incident supersonic flow were taken equal to 100 kPa and 293.15 K, respectively. Special calculations with the expansion of the computational domain showed that the adopted boundary conditions did not affect the solution of the problem.

In FIG. Figure 2 shows the calculated dependences of the specific impulse of a ramjet with continuous detonation combustion of a mixture of hydrogen and air in the range of flight Mach numbers from 1.7 to 2.2.

The calculation showed that in the considered scheme the specific impulse exceeds 2700 s, and the proposed method and device in the entire studied range of flight Mach numbers provide a positive effective thrust. This means that an aircraft equipped with a ramjet engine with continuous detonation combustion can carry out autonomous supersonic flight, starting with a Mach number exceeding 1.7, which is significantly less than that of the prototype device. This conclusion is generally confirmed by the experimental data obtained in the work (Frolov SM, Zvegintsev VI, Ivanov VS, Aksenov VS, Shamshin IO, Vnuchkov DA, Nalivaichenko DG, Berlin AA, Fomin VM Wind tunnel tests of a hydrogen-fueled detonation ramjet model at approach air stream Mach numbers from 4 to 8. International Journal of Hydrogen Energy, 2017, Vol. 42, pp. 25401-25413). In this work, it was shown that at an incident air velocity in the inlet section of a supersonic air intake approximately corresponding to the local Mach number from 1 to 1.75, it is possible to realize continuous detonation combustion of hydrogen in the annular combustion chamber (see Fig. 10 in the cited article).

Claims (7)

1. A method of organizing a working process in a ramjet engine with continuous detonation combustion, including accelerating an aircraft to supersonic speed, ensuring the start of an autonomous flight of an aircraft with such an engine, partial braking of the incident supersonic air flow in oblique shock waves and in the near-boundary boundary layer before entering the annular combustion chamber, the continuous supply of fuel to the mixing zone with air, the formation of detonation-capable of a mixture of fuel and air, initiating continuous detonation combustion of the fuel mixture, outflow of detonation products from the annular combustion chamber through an annular nozzle with supersonic speed with the formation of a jet stream and the creation of reactive thrust, characterized in that the incident supersonic air flow is first partially inhibited in oblique jumps compaction in the parietal boundary layer, and then accelerated in a fan of rarefaction waves with a partial restoration of the parameters of the incident supersonic air flow and enters the ramjet in the form of a slightly braked supersonic air flow, while one part of the air is directed into the annular combustion chamber to form a detonation-capable mixture of fuel and air at such a distance from the inlet section of the annular combustion chamber so that it is continuously detonation combustion in the annular combustion chamber did not lead to the displacement of the mixture of fuel and air, as well as hot detonation products from the annular combustion chamber through its inlet section, and The second part, including the wall boundary layer, is bypassed by the annular combustion chamber to provide cooling of the walls of the annular combustion chamber and to prevent the gas-dynamic effect of continuous detonation combustion of the fuel-air mixture in the annular combustion chamber on the flow of weakly inhibited supersonic air flow at the inlet to the direct-flow air -jet engine. 2. A device for implementing a method of organizing a working process in a ramjet with continuously detonating combustion, including a supersonic air intake, a central body with front and rear cones, an annular combustion chamber with a belt of fuel nozzles and an annular nozzle, characterized in that the central the body contains an additional section between the front and rear cones, and the supersonic air intake communicates with the annular combustion chamber and with the gas-dynamic insulator, in the form of a bypass channel and located between the annular combustion chamber and the outer wall of the rear cone of the central body, the front edge of the wall separating the annular combustion chamber from the gas-dynamic insulator, and shifted deep into the supersonic air intake, and the rear edge of the wall separating the annular combustion chamber from the gas-dynamic insulator, offset deep into the annular nozzle. 3. The device according to p. 2, characterized in that the additional portion of the central body may have a conical, cylindrical or other shape, providing an uninterrupted flow of supersonic air flow. 4. The device according to claim 2, characterized in that the supersonic air intake can be made according to the internal compression scheme by profiling its internal path or according to the compression scheme in the attached oblique shock waves. 5. The device according to p. 2, characterized in that the fuel nozzles in the annular combustion chamber are located at such a distance from the input section of the annular combustion chamber so that continuous detonation combustion in the annular combustion chamber does not lead to the displacement of the mixture of fuel and air, and hot detonation products from the annular combustion chamber through its inlet section. 6. The device according to p. 2, characterized in that the rear cone of the axisymmetric central body can be truncated and have the form of a straight cone or cone with a profiled side surface. 7. The device according to p. 2, characterized in that the annular nozzle is preferably made in the form of an expanding annular channel to ensure complete expansion of the detonation products coming from the annular combustion chamber and the air flow coming from the gas-dynamic insulator to atmospheric pressure without shock waves .

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