RU2591391C1 - Vortex low-thrust rocket engine on gas fuel - Google Patents

Abstract

FIELD: missile.

SUBSTANCE: invention relates to rocket engineering and can be used in development of rocket engines operating on gaseous fuel mixture components. Vortex low-thrust rocket engine on gas fuel comprises combustion chamber with nozzle and tangential swirlers for supply of fuel components. In combustion chamber on side of its front bottom there are internal and external swirlers made coaxially and separated by cylindrical barrel. Swirling chamber of inner swirler with radius R₁ is inner surface of shell, swirling chamber of outer swirler with radius R₂ is inner surface of combustion chamber. Tangential channels to feed fuel components to internal and external swirlers are directed oppositely. Height of cylindrical barrel is defined by relationship h = (0.4÷0.6)R₂, and combustion chamber has a narrowing of radius R₂ to radius R₁ at distance L = (1÷1.5)R₂ from front bottom. Ratio of radii of swirling chambers of outer and inner swirlers is determined by solving an algebraic equation.

EFFECT: invention provides operation with any gaseous fuel compositions and high energy-traction characteristics due to increased completeness of combustion at intensive mixing fuel and oxidiser in opposite swirled flows, as well as compensation of reactive force causing undesirable rotation of engine.

1 cl, 4 dwg

Description

The invention relates to the field of rocket technology and can be used in the development of rocket engines running on gaseous components of the fuel mixture.

Small thrust rocket propulsion systems (up to 1,500 N) are subsystems of spacecraft flight control systems — their executive bodies [1]. Currently, liquid propellant rocket engines (LRE) and solid propellant rocket engines (solid propellant rocket engines) are mainly used as control propulsion systems [2]. Gaseous-propellant rocket engines can be used to create special propulsion systems, in particular when creating low-thrust rocket engines to remove spent launch vehicle stages from occupied orbits using gasified liquid components of the guaranteed liquid fuel reserve rocket engine as gaseous fuel [3].

A known method of organizing a workflow in a combustion chamber of a rocket engine [4], in which self-igniting fuel components are used. At the same time, they are fed into the combustion chamber through tangential entries into the corresponding coaxial twisting chambers of the two-component centrifugal nozzle. The twisting of the components contributes to a more complete mixing of the components of the fuel mixture and provides thermal protection of the combustion chamber.

The patent [5] proposes the use of a prechamber (prechamber) for swirling and mixing the fuel mixture by supplying gaseous components of the fuel with the help of a screw.

Known vortex rocket engine, in which to organize the processes of mixture formation and combustion of fuel components using their twist [6]. In this case, the main part of the components is supplied from the nozzle block side through nozzles located evenly tangentially to the circle at an angle of 60 ° to the surface of the roof of the combustion chamber. The engine is equipped with deflecting blades to compensate for the jet rotational movement of the chamber, mounted on a cone, which is paired with the neck of the chamber.

To increase the specific impulse of engine thrust, the patent [7] proposes the supply of non-combustible components into the chamber by means of tangential inlets of a gaseous oxidizer and jet nozzles of liquid fuel. In this case, the resulting swirling flow of a gaseous oxidizer and a spray jet of fuel are mixed and fed into the combustion chamber.

A known combustion chamber of a liquid propulsion thruster [8], in which the components of the fuel mixture are sprayed by coaxial centrifugal nozzles with the opposite direction of twist. In this case, the manifold of the outer nozzle is in communication with the jet nozzles evenly spaced around the circumference.

Closest to the technical nature of the claimed invention is a rocket engine gaseous fuel [9]. Gaseous fuel and oxidizing agent, pre-mixed in a prechamber with helium gas and aluminum powder with an average particle size of not more than 10 microns, enter through the tangential inlets into the combustion chamber from the nozzle side of the engine cover.

The disadvantage of this propulsion system is the inability to use self-igniting fuel components, the ignition of which can occur in the prechamber, as well as fuels with significantly different phase transition temperatures due to the formation of condensate during mixing.

The technical result of the present invention is the development of a small thrust rocket engine on gaseous fuel, which ensures reliable operation with any gaseous fuel compositions, high energy traction characteristics by increasing the completeness of combustion with vigorous mixing of the fuel and oxidizing agent in counter swirling flows, and also compensating for reactive forces causing undesirable engine rotation.

The technical result of the invention is achieved by the fact that a vortex rocket engine of low thrust on gaseous fuel is developed, including a combustion chamber with a nozzle and tangential swirls for supplying fuel components to the combustion chamber. In the combustion chamber from the side of its front bottom there are two coaxially made swirls for separate supply of fuel components (internal and external), separated by a cylindrical glass. The twisting chamber of the inner swirl is the inner surface of the cylindrical cup, the twisting chamber of the outer swirl is the inner surface of the combustion chamber, the tangential channels for supplying fuel components in the inner and outer swirls are opposite. The combustion chamber is constricted from a radius R ₂ to a radius R ₁ located at a distance L = (1 ÷ 1.5) R ₂ from the front bottom of the combustion chamber, the height of the cylindrical cup is determined by the ratio

h = (0.4 ÷ 0.6) R ₂ ,

and the ratio of the radii of the swirl chambers of the external and internal swirls is determined from the solution of the equation

- константа, определяемая конкретной топливной композицией (горючее и окислитель); ρ₁, ρ₂, - плотности компонентов топлива, подаваемых во внутренний и внешний завихрители, соответственно, кг/м³; G₁, G₂ - массовые секундные расходы компонентов топлива, подаваемых во внутренний и внешний завихрители, соответственно, кг/с.where h is the height of the cylindrical glass; L is the distance from the front bottom of the combustion chamber to its narrowing; R ₁ , R ₂ - the radii of the swirl chambers of the internal and external swirlers; r = R ₂ / R ₁ is the ratio of the radii of the swirl chambers of swirls; $$C=0.75\frac{G_2}{G_{\mathrm{one}}}\sqrt{\frac{{\rho }_{\mathrm{one}}}{{\rho }_2}}$$

- a constant determined by a specific fuel composition (fuel and oxidizer); ρ ₁ , ρ ₂ , are the densities of the fuel components supplied to the internal and external swirlers, respectively, kg / m ³ ; G ₁ , G ₂ - mass second consumption of fuel components supplied to the internal and external swirlers, respectively, kg / s

The achievement of the positive effect of the invention is provided by the following factors.

1. Separate supply of fuel components to the combustion chamber and the absence of nozzles and atomizers allows the use of any gaseous fuel components, including those containing condensed inclusions (particles), which ensures reliable engine operation.

2. The wall height of the cylindrical glass h = (0.4 ÷ 0.6) R ₂ separating the external and internal swirls allows one to organize the formation of a stable vortex in both the external and internal swirls. A calculated estimation of the characteristics of the swirling gas flow in a cylindrical channel showed that when the wall height of the cylindrical glass is less than 0.4R ₂ , the stability of the internal vortex and the formation of the axial component of the velocity of motion are not ensured. When the height of the wall of the cylindrical glass is greater than 0.6R ₂ , the tangential velocity component decreases due to braking of the vortex on the wall of the swirl chamber and the weight and dimensions of the combustion chamber increase.

3. The narrowing of the combustion chamber to a radius R ₁ equal to the radius of the chamber of the internal swirl ensures the displacement of the external vortex towards the internal one, which, under the action of centrifugal forces, after leaving the cylindrical glass separating the swirls, moves towards the external vortex. The superposition of vortices on one another leads to inhibition of the tangential component of the velocity of the fuel components and their intensive mixing. When the narrowing of the combustion chamber is smaller than the radius of the chamber of the internal swirl R ₁ , complete overlap of the vortices does not occur and the mixing conditions of the fuel components deteriorate. When the combustion chamber narrows more than R ₁ , it is possible to "lock" the external vortex in the space between the walls of the combustion chamber and the cylindrical cup of the internal swirl, which worsens the mixing conditions of the fuel components.

4. The location of the narrowing of the combustion chamber at a distance L = (1 ÷ 1.5) R ₂ from the front bottom of the combustion chamber provides the greatest efficiency of the engine. With an increase in this distance, the dimensions and mass of the engine increase, and with its decrease, the interaction zone of vortices decreases and, therefore, the mixing conditions of the fuel components deteriorate.

5. Counter-swirling gas vortices with the same intensity provides complete braking of the tangential component of the vortex velocity and compensates for the reactive force arising from the tangential introduction of fuel components, which can cause the engine to rotate.

The tangential swirl swirl intensity is determined by the dimensionless Higir-Baer parameter [10]:

where R is the radius of the swirl chamber;

ρ is the gas density;

u is the axial component of velocity;

w is the tangential component of velocity;

r is the current radius.

We define the Higier-Baire parameter for the internal swirler. Suppose that the density ρ ₁ and the axial component of the gas velocity u _{1 are} constant, and the radial distribution of the tangential component of the velocity w (r) corresponds to the law of a solid [10]

where w ₁ is the velocity of the gas entering the chamber of the internal swirler from the tangential channels at a radius of R ₁ .

Carrying out integration (1) taking into account (2), for the internal swirler we obtain the formula for calculating Ф ₁ :

For the external swirl, integration is carried out from the radius of the internal swirl R ₁ (the wall thickness of the cylindrical glass can be neglected) to the radius of the external swirl R ₂ . In this case, gas rotation occurs in the annular channel and a change in the tangential velocity of the gas along the radius can be neglected (w ₂ = const). For an external swirler, the Higier-Baer parameter is:

Expressing the gas velocity through its flow rate G = ρuS (ρ is the gas density, u is the gas velocity; S is the cross-sectional area), we obtain the equations for the components of the velocity vector in the internal and external swirlers:

where S ₁ , S ₂ - the area of the input tangential channels for the internal and external swirls.

The spin intensity for both swirlers should be the same, therefore, substituting (5) in (3) and (4) and equating the Higir-Baer parameters (Ф ₁ = Ф ₂ ), we obtain:

The mass flow rate G is related to the pressure drop Δp on the swirler input channels by the ratio [11]:

where φ is the flow coefficient;

S is the total area of the input tangential channels.

Assuming that the supply of fuel components to the engine is carried out at the same pressure drop (Δp ₁ = Δp ₂ ), and the flow coefficients of the input channels are equal (φ ₁ = φ ₂ ), from the equation (7) we can obtain the ratio of the total areas of the input channels of internal swirls:

From relations (6) and (8), we obtain an equation for determining the ratio of the radii of the swirl chambers, which provides the same intensity of swirling of the fuel components at a given pressure drop:

Equation (9) is converted to a cubic equation with respect to r = R ₂ / R ₁ .

- константа для конкретной топливной композиции.Where $$C=0.75\frac{G_2}{G_{\mathrm{one}}}\sqrt{\frac{{\rho }_{\mathrm{one}}}{{\rho }_2}}$$

- constant for a specific fuel composition.

Choosing the values of the radius of the swirl chamber and the total area of the tangential channels for one fuel component, the corresponding values for the swirl chamber of the second component are determined using formulas (8), (10), which ensure the same intensity of the twisting of the fuel components.

The invention is illustrated by the scheme of the vortex rocket engine of low thrust on gaseous fuels that implements the proposed invention (Fig. 1). The rocket engine comprises a combustion chamber 1 and a nozzle 2. In FIG. 2 and FIG. 3 shows sections of chambers of inner 3 of radius R ₁ and outer 4 of radius R _{2 of} tangential swirls arranged coaxially. Swirlers 3 and 4 are separated by a cylindrical glass 5 of height h = (0.4 ÷ 0.6) R ₂ . On the front bottom 6 of the combustion chamber 1 are one or more pyrotechnic igniters (not shown in FIG. 1). One of the gaseous components of the fuel is fed through a gas pipeline 7 to the manifold 8 and through the tangential channels 9 into the swirl chamber of the internal swirl 3 (Fig. 2). The second component of the fuel is supplied through a gas pipeline 10 to the collector 11 and through the tangential channels 12 into the twisting chamber of the external swirl 4 (Fig. 3). The rotation of the gas in the swirls is directed oppositely to each other. To ensure uniform swirling, several tangential channels are used to supply fuel components. The combustion chamber 1 at a distance L = (1 ÷ 1.5) R ₂ from the front bottom narrows to a radius R ₁ (Fig. 1). The ratio of the radii of the swirl chambers of the swirls 3 and 4 should satisfy equation (10), and the ratio of the total areas of the tangential channels 9 and 12 to the relation (8).

The engine operates as follows. Gaseous fuel and oxidizing agent are supplied under pressure to the combustion chamber directly from the tanks or from the gasification device (not shown in Fig. 1). For example, an oxidizing agent is supplied through a gas pipeline 10 to a collector 11 and through tangential channels 12 to a swirl chamber of an external swirl 4, and fuel through a gas pipeline 7 enters a collector 8 and through tangential channels 9 to a swirl chamber of an internal swirl 3. A gaseous oxidizer moving along the chamber wall combustion 1, is shifted to the axis of the chamber due to the narrowing of the chamber. Gaseous fuel, rotating in the chamber of the internal swirler 3, moves towards the nozzle 2 along the wall of the cylindrical cup 5 and, leaving it, under the action of centrifugal forces moves to the wall of the combustion chamber 1, where it interacts with the oxidizer vortex. Since the directions of rotation of the oxidizing agent and the fuel are opposite, and the swirl intensities are the same, the tangential component of the velocity of the vortices is decelerated and intensively mixed. The mixed fuel components are either self-igniting (for self-igniting components) or ignited by a pyrotechnic igniter. The engine goes to stationary operation. If necessary, a multiple start mode can be implemented by interrupting the supply of fuel components and re-starting it with subsequent ignition of the fuel mixture.

Implementation example

Let us cite as an example the calculation of a low-thrust vortex rocket engine using gaseous fuel with a thrust of P = 500 N operating on gaseous methane and oxygen. Methane has a molecular weight of µ = 16 g / mol and a density of ρ = 0.668 kg / m ³ under normal conditions. For oxygen, these values are μ = 64 g / mol and ρ = 1.331 kg / m ³ .

Chemical reaction of methane combustion in oxygen

CH ₄ + 2O ₂ = CO ₂ + 2H ₂ O

shows that for one gram-molecule of methane, 2 gram-molecules of oxygen are required, or in the mass ratio - for 16 g of methane 64 g of oxygen are required, i.e. the stoichiometric coefficient is 4. The magnitude of the engine thrust can be estimated from the relation [2]:

P = G · I,

where G is the mass second fuel consumption; I is the specific impulse of thrust. For the fuel mixture under consideration (methane + oxygen), the specific thrust impulse calculated using the TERRA software system is I = 388 m / s. Then the fuel consumption for an engine with a thrust of P = 500 N is equal to:

For a stoichiometric mixture, the oxidizer consumption will be 1.032 kg / s, and for fuel - 0.258 kg / s.

Option 1

The oxidizing agent is supplied to the external swirl, and the fuel to the internal. In this case, G ₁ = 0.258 kg / s, ρ ₁ = 0.668 kg / m ³ , and G ₂ = 1.032 kg / s, ρ ₂ = 1.331 kg / m ³ .

Substituting these values in the formula (10), we define the constant C:

From the graphical solution of equation (10) we find the value of r ₁ (Fig. 4):

We take the radius of the combustion chamber R ₂ = 40 mm, then the value of R ₁ = 24 mm.

For a given pressure drop on the channels 9 and 12 of the tangential swirls Δp = 1 MPa, using the formula (7), we determine the total area of the input channels (we assume that the flow coefficient is equal to φ ₁ = 0.85):

Let the internal swirl have n ₁ = 6 tangential channels, then the diameter of one channel will be equal to:

From relation (8) we determine the total area of the openings of the tangential channels for the external swirl:

Let the external swirl have n ₁ = 12 tangential channels, then the diameter of one channel will be equal to:

The gaseous fuel thrust vortex rocket engine calculated in accordance with the invention has an external swirl with a swirl chamber with a diameter of 80 mm and 12 tangential channels with a diameter of 8.9 mm, an internal swirl with a chamber with a diameter of 48 mm and 6 tangential channels with a diameter of 7.5 mm. The swirlers are separated by a cylindrical glass with a height of h = 0.5R ₂ = 20 mm, and the combustion chamber itself narrows to a diameter of D = 2R ₁ = 48 mm at a distance equal to l = 1.2R ₂ = 48 mm from the rear bottom.

Option 2

Fuel is supplied to the external swirl, and the oxidizer to the internal swirl. In this case, G ₁ = 1.032 kg / s, ρ ₁ = 1.331 kg / m ³ , a G ₂ = 0.258 kg / s, ρ ₂ = 0.668 kg / m ³ .

Substituting these values in the formula (10), we define the constant C:

From the graphical solution of equation (10) we find the value of r ₂ (Fig. 4):

Choosing the same value of R ₂ = 40 mm for the swirl chamber of the external swirl, for R ₁ we get R ₁ = 36.7 mm.

Since the costs of the fuel components do not change, the total area of the supply openings, their number and diameters will remain the same. The engine calculated in accordance with the invention, when the oxidant is fed into the internal swirl, has an external swirl with a chamber of 80 mm diameter and 6 tangential channels with a diameter of 7.5 mm, an internal swirl with a chamber of 73.4 mm diameter and 12 tangential channels with a diameter of 8.9 mm. The swirlers are separated by a cylindrical glass with a height of h = 0.5R ₂ = 20 mm, and the combustion chamber itself narrows to a diameter of D = 2R ₁ = 73 mm at a distance equal to l = 1.2R ₂ = 48 mm from the rear bottom.

Thus, the inventive short-thrust vortex rocket engine using gaseous fuel ensures the achievement of the technical result of the invention — reliable operation with any gaseous fuel compositions, high energy traction characteristics due to increased completeness of combustion with vigorous mixing of the fuel and oxidizer in counter-swirling flows, as well as reactive force compensation causing unwanted engine rotation.

LITERATURE

1. Grishin S.D., Kokorin V.V., Kharlamov N.P. The theoretical basis for the creation of propulsion systems for controlling spacecraft. M .: Mechanical Engineering, 1985 .-- 192 p.

2. Melkumov TM, Melik-Pashaev NI, Chistyakov P. G., Shiukov A. G. Rocket engines. M .: Mechanical Engineering, 1976 .-- 400 p.

3. Belokonov I.V., Kruglov G.E., Trushlyakov V.I., Yudintsev V.V. Evaluation of the possibility of controlled descent from the orbit of the upper stage of the Soyuz launch vehicle through the use of fuel residues in tanks. // All-Russian scientific and technical conference "Actual problems of rocket and space technology and its role in the sustainable socio-economic development of society", dedicated to the 50th anniversary of the founding of the Central Design Bureau and the 90th anniversary of the birth of D.I. Kozlova. Samara 2009 .-- S. 68-72.

4. RF patent No. 2192556 C2, IPC F02K 9/56, F02K 9/52. The method of organizing the working process in the chamber of a liquid propulsion thruster / Kazankin F.A., Kutuev R.Kh., Larin EG, Mezenin PB .; publ. November 10, 2002

5. RF patent No. 2183761 C2, IPC F02K 9/62, F02K 9/95. A liquid propulsion thruster and a method for starting a fluid thruster / Vesnovatov AG, Barsukov OA; publ. 06/20/2002

6. RF patent No. 2300007 C1, IPC F02K 9/62. Vortex rocket engine / Timoshenko I.K .; publ. May 27, 2007

7. RF patent No. 2397355, IPC F02K 9/62, F02K 9/95. A method of organizing the working process of a small thrust rocket engine / Kutuev R.Kh .; publ. 08/20/2010 r.

8. RF patent №2217620, IPC F02K 9/62, F02K 9/52. Chamber of a liquid propellant rocket engine of small thrust / Ivanov V.N .; publ. November 27, 2003

9. RF patent No. 2488712 C2, IPC F02K 9/62. A method of organizing a work process in a space propulsion system using gaseous fuel / Arkhipov V.A., Borisov B.V., Zhukov A.S., Bondarchuk S.S., Kudentsov V.Yu., Trushlyakov V.I .; publ. 07/27/2013

10. Gupta A., Lilly D., Cyred N. Swirling flows. - M .: Mir, 1987 .-- 588 p.

11. The Kremlin P.P. Flow meters and quantity counters. - L: Engineering, 1989 .-- 701 p.

Claims (1)

A gaseous fuel thrust vortex rocket engine comprising a combustion chamber with a nozzle and tangential swirls for supplying fuel components to the combustion chamber, characterized in that two coaxially made swirls for separately supplying fuel components are located in the combustion chamber from the side of its front bottom (internal and external), separated by a cylindrical cup, while the twist chamber of the inner swirl is the inner surface of the cylindrical cup, the twist chamber the external swirl is the inner surface of the combustion chamber, the tangential channels of the supply of fuel components in the inner and outer swirls are directed opposite, the combustion chamber is made narrower from a radius R ₂ to a radius R ₁ located at a distance L = (1 ÷ 1.5) R ₂ from the front bottom combustion chamber, while the height of the cylindrical glass is determined by the ratio
h = (0.4 ÷ 0.6) R ₂ ,
and the ratio of the radii of the swirl chambers of the external and internal swirls is determined from the solution of the equation

where h is the height of the cylindrical glass;
L is the distance from the front bottom of the combustion chamber to its narrowing;
R ₁ is the radius of the swirl chamber of the internal swirl;
R ₂ is the radius of the swirl chamber of the external swirl;
r = R ₂ / R ₁ is the ratio of the radii of the swirl chambers of swirls;

- a constant determined by a specific fuel composition (fuel and oxidizer);
ρ ₁ , ρ ₂ - the density of the components of the fuel supplied to the internal and external swirlers, respectively, kg / m ³ ;
G ₁ , G ₂ - mass second consumption of fuel components supplied to the internal and external swirlers, respectively, kg / s

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