Scramjet engine combustion chamber thrust control method during acceleration process

Abstract

The application discloses a combustion chamber thrust control method in the acceleration process of a scramjet engine, and adopts a gradually expanding series connection cavity combustion chamber, fuel injection holes are arranged on the upstream of two cavities, the first cavity is a pre-combustion zone, and the second cavity is a main combustion zone; the thrust control method is as follows: when the combustion chamber inlet Mach number is low, the combustion mode of the pre-combustion zone is controlled to be kept in a weak subsonic combustion mode, and the combustion mode of the main combustion zone is controlled to be kept in a strong subsonic combustion mode; when the combustion chamber inlet Mach number is high, the combustion mode of the pre-combustion zone is controlled to be kept in a supersonic combustion mode, and the combustion mode of the main combustion zone is controlled to be kept in a strong subsonic combustion mode. The application is applied to the field of scramjet engines, can guarantee that the combustion chamber can generate higher thrust in a wide flight Mach number range, simultaneously avoids thermal choking, and the thrust mutation caused by combustion mode conversion is smaller, thrust control is more stable, and smooth transition of thrust during combustion mode conversion can be realized.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of scramjet, in particular to a method for controlling the thrust of a combustion chamber in the acceleration process of a scramjet. BACKGROUND

[0002] A hypersonic vehicle using a scramjet as a propulsion device will experience an acceleration process in the near space flight. In this process, the Mach number and the flight altitude of the vehicle are constantly increasing, the inflow conditions of the combustion chamber are constantly changing, and the combustion mode also changes accordingly. Studies have shown that the mode change will cause a sudden change in thrust performance, which will adversely affect the thrust control of the hypersonic vehicle.

[0003] Generally, the hypersonic vehicle uses the method of adjusting the equivalence ratio of fuel to control the thrust size of the combustion chamber. During the acceleration process of the hypersonic vehicle, the inlet Mach number of the combustion chamber is constantly changing, and the existing thrust control method mainly focuses on the research under single inflow condition, which has a large difference from the combustion chamber environment in the acceleration process. Studies have shown that under the same inflow Mach number condition, if the combustion mode does not change, the thrust coefficient is approximately linearly related to the equivalence ratio of fuel. On the basis of this understanding, by modeling the fuel supply and engine thrust performance, a better thrust control scheme can be obtained by using a multi-objective optimization method. However, in the process of adjusting the equivalence ratio of fuel, the combustion mode in the supersonic combustion chamber is easily changed due to the change of the inflow condition, which will cause the combustion chamber thrust and the equivalence ratio of fuel to present a strong nonlinear relationship, and further cause the combustion chamber thrust performance to appear a sudden change, which makes the original fuel injection scheme of the combustion chamber difficult to apply. At the same time, this change has a certain hysteresis and is usually difficult to predict, so it is very difficult to achieve precise control of the combustion mode of the combustion chamber under variable inflow conditions through fuel injection control.

[0004] In the real flight process, the combustion mode will be affected by many factors such as inflow Mach number, equivalence ratio, combustion chamber configuration, etc. Therefore, in the acceleration process of the hypersonic vehicle, when adjusting the equivalence ratio to achieve thrust performance control, the combustion mode is likely to change unpredictably, causing a sudden change in the combustion chamber thrust, and further affecting the thrust control of the hypersonic vehicle. Therefore, it is valuable to develop a fuel injection scheme that has more stable thrust control effect and smoother thrust change in the acceleration process. SUMMARY

[0005] In view of the deficiencies in the prior art, the present application provides a method for controlling the thrust of a combustion chamber in the acceleration process of a scramjet, which adopts a two-stage fuel injection scheme to ensure that the combustion chamber can generate higher thrust in a wide range of flight Mach numbers, and to achieve smooth transition of the thrust when the combustion mode changes.

[0006] To achieve the above object, the present application provides a method for controlling the thrust of a combustion chamber in the acceleration process of a scramjet engine, which adopts a series of gradually expanding cavity combustion chambers, fuel injection holes are arranged upstream of both cavities, the first cavity is a pre-combustion zone, and the second cavity is a main combustion zone.

[0007] The thrust control method is characterized in that:

[0008] When the Mach number of the combustion chamber inlet is low, the combustion mode of the pre-combustion zone is controlled to remain in a weak subsonic mode, and the combustion mode of the main combustion zone is controlled to remain in a strong subsonic mode.

[0009] When the Mach number of the combustion chamber inlet is high, the combustion mode of the pre-combustion zone is controlled to remain in a super-sonic mode, and the combustion mode of the main combustion zone is controlled to remain in a strong subsonic mode.

[0010] In one embodiment, the thrust control method is characterized in that:

[0011] When the Mach number of the combustion chamber inlet is lower than the critical Mach number of the combustion mode conversion of the pre-combustion zone, the equivalence ratio of the pre-combustion zone is controlled to make the combustion mode remain in a weak subsonic mode, and the equivalence ratio of the main combustion zone is controlled to make the combustion mode remain in a strong subsonic mode.

[0012] When the Mach number of the combustion chamber inlet is higher than the critical Mach number of the combustion mode conversion of the pre-combustion zone, the equivalence ratio of the pre-combustion zone is controlled to make the combustion mode remain in a super-sonic mode, and the equivalence ratio of the main combustion zone is controlled to make the combustion mode remain in a strong subsonic mode.

[0013] In one embodiment, the process of controlling the equivalence ratio of the pre-combustion zone is characterized in that:

[0014] The energy and mass conservation equations are established in the pre-combustion zone, which are:

[0015]

[0016] wherein, is the mass flow rate of the combustion chamber inlet, k1 is the air-fuel ratio of the pre-combustion zone, P 1c is the pressure of the pre-combustion zone core flow, R is the gas constant, T 1c is the temperature of the pre-combustion zone core flow, T wall is the near-wall temperature of the pre-combustion zone, h1 is the height of the pre-combustion zone, W is the width of the isolation section, γ is the specific heat ratio, h 1c is the height of the pre-combustion zone core flow, Ma 1c is the Mach number of the pre-combustion zone core flow, C p is the isobaric heat capacity, is the pre-combustion zone mass average temperature except the cavity, T is the temperature of the isolation section, ρ1 is the density of the pre-combustion zone, V1 is the volume of the pre-combustion zone, V cavity is the cavity volume, TMax is the maximum temperature in the pre-combustion cavity, η1 is the combustion efficiency, E fuel is the fuel heat value, ER1 is the equivalence ratio of the pre-combustion cavity, k0 is the equivalence fuel-air ratio, L1 is the length of the pre-combustion cavity, u1 is the average velocity of the jet wake gas in the pre-combustion cavity;

[0017] By solving the energy and mass conservation equations of the pre-combustion cavity, the length L1 of the pre-combustion cavity under the condition of a given equivalence ratio is obtained. In the process of controlling the increase of the equivalence ratio of the pre-combustion cavity, when the length L1 of the pre-combustion cavity covers the pre-combustion cavity and the position of 1 / 2 of the distance from the pre-combustion cavity to the main combustion cavity, it is considered that the equivalence ratio ER1 of the current pre-combustion cavity reaches the required value.

[0018] In one of the embodiments, in the process of controlling the equivalence ratio of the main combustion cavity to maintain the combustion mode in the strong subsonic mode, the thermal choking caused by the too high equivalence ratio of the main combustion cavity is avoided to prevent the decline of the thrust performance.

[0019] In one of the embodiments, the process of controlling the equivalence ratio of the main combustion cavity is as follows:

[0020] The energy and mass conservation equations of the main combustion cavity are established, which are as follows:

[0021]

[0022] wherein, is the mass flow rate at the inlet of the combustion chamber, k1 is the air-fuel ratio of the pre-combustion cavity, k2 is the air-fuel ratio of the main combustion cavity, P 2c is the pressure of the core flow of the main combustion cavity, R is the gas constant, T 2c is the temperature of the core flow of the main combustion cavity, A 2c is the cross-sectional area of the core flow of the main combustion cavity, Ma 2c is the Mach number of the core flow of the main combustion cavity, γ is the specific heat ratio, P2 is the pressure of the main combustion cavity, T2 is the temperature of the main combustion cavity, W is the width of the isolation section, h2 is the height of the main combustion cavity, C p is the isobaric heat capacity, T is the temperature of the isolation section, ρ2 is the density of the main combustion cavity, V2 is the volume of the main combustion cavity, η1 is the combustion efficiency of the pre-combustion cavity, ER1 is the equivalence ratio of the pre-combustion cavity, η2 is the combustion efficiency of the main combustion cavity, ER2 is the equivalence ratio of the main combustion cavity, k0 is the equivalence fuel-air ratio, E fuel is the fuel heat value, L2 is the length of the main combustion cavity, u2 is the average velocity of the jet wake gas in the main combustion cavity;

[0023] By solving the energy and mass conservation equations of the main combustion cavity, h1 and p1 under the condition of a given equivalence ratio of the main combustion cavity are obtained. In the process of controlling the increase of the equivalence ratio of the main combustion cavity, h2≤H+x2·tanα is maintained, wherein H is the half height of the combustion chamber, x2 is the horizontal distance from the midpoint of the main combustion cavity to the expansion corner of the combustion chamber, and α is the expansion angle of the combustion chamber.

[0024] Compared with the prior art, the present application has the following beneficial technical effects:

[0025] 1. The scramjet combustor of the present application can produce higher thrust while avoiding thermal choking by using the injection scheme;

[0026] 2. The combustor mode conversion of the present application results in smaller thrust mutation and more stable thrust control;

[0027] 3. The present application is suitable for hypersonic flight in a wide range of flight Mach numbers. BRIEF DESCRIPTION OF DRAWINGS

[0028] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or prior art description. Obviously, the drawings in the following description only constitute some embodiments of the present application, and for those skilled in the art, other drawings can also be obtained from the structures shown in the drawings without creative labor.

[0029] Figure 1 Structure diagram of the gradually expanding series cavity combustor in the embodiment of the present application;

[0030] Figure 2 Combustor region division diagram in the embodiment of the present application;

[0031] Figure 3 Combustor working mode diagram at a lower Mach number in the embodiment of the present application;

[0032] Figure 4 Combustor working mode diagram at a higher Mach number in the embodiment of the present application.

[0033] The implementation, functional features and advantages of the present application will be further described with reference to the embodiments and the accompanying drawings. DETAILED DESCRIPTION

[0034] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments only constitute some of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor fall within the protection scope of the present application.

[0035] In addition, the technical solutions of each embodiment of the present application can be combined with each other, but it must be based on the fact that a person skilled in the art can realize it. When the combination of technical solutions appears contradictory or unachievable, it should be considered that the combination of technical solutions does not exist and is not within the protection scope of the present application.

[0036] The embodiment discloses a combustion chamber thrust control method in a scramjet engine acceleration process, adopts a gradually expanding series of concave cavity combustion chambers to carry out two-stage fuel injection in a hypersonic vehicle acceleration process, can ensure that the combustion chamber can generate higher thrust in a wide flight Mach number range, and can realize smooth transition of thrust in combustion mode conversion.

[0037] Reference Figure 1 In the embodiment, the gradually expanding series of concave cavity combustion chambers are provided with fuel injection holes upstream of two concave cavities, namely Jet1 and Jet2 in Figure 1 The combustion zone of the gradually expanding series of concave cavity combustion chambers is mainly divided into two parts, namely a pre-combustion zone and a main combustion zone. The pre-combustion zone is located at the position of the first concave cavity, and the main combustion zone is located at the position of the second concave cavity. Among them, the combustion mode of the pre-combustion zone mainly depends on the incoming flow Mach number; and for the main combustion zone, due to the deceleration effect of the pre-combustion zone combustion heat release, the combustion mode of the main combustion zone mainly depends on the equivalence ratio, as shown in Table 1.

[0038] Table 1 Main influencing factors of combustion mode of main combustion zone in front of series of concave cavity combustion chamber

[0039]

[0040] Based on this, the thrust control method adopted in the embodiment is specifically:

[0041] When the combustion chamber inlet Mach number is lower than the pre-combustion zone combustion mode conversion critical Mach number, the equivalence ratio of the pre-combustion zone is controlled to keep the combustion mode in a weak subsonic mode, and the equivalence ratio of the main combustion zone is controlled to maintain the combustion mode in a strong subsonic mode;

[0042] When the combustion chamber inlet Mach number is higher than the pre-combustion zone combustion mode conversion critical Mach number, the equivalence ratio of the pre-combustion zone is controlled to keep the combustion mode in a super-sonic mode, and the equivalence ratio of the main combustion zone is controlled to maintain the combustion mode in a strong subsonic mode.

[0043] Among them, the acquisition process of the pre-combustion zone combustion mode conversion critical Mach number is:

[0044] First, the gradually expanding series of concave cavity combustion chambers are divided into five main regions, namely isolation section, transition zone, pre-combustion zone, main combustion zone and core flow, as shown in Figure 2 For different regions, the physical property parameters have great differences, and for the same region, the physical property parameters have the characteristics of uniform distribution or linear change.

[0045] According to the core flow model of the isolation section, the Mach number Ma before the transition zone can be calculated by the following formula:

[0046]

[0047]

[0048]

[0049] where γ is the specific heat ratio, γ in is the specific heat ratio at the combustor inlet, Ma in is the Mach number at the combustor inlet, π is the pressure ratio of the isolator to the combustor inlet, A is the cross-sectional area of the combustor at the end of the isolator shock train, A in is the cross-sectional area of the combustor at the inlet, A c is the core flow area at the end of the isolator shock train;

[0050] The length of the isolator shock train L0 can be calculated by the following formula, which is:

[0051]

[0052] where D is the equivalent diameter of the isolator, c1 and c2 are 50 and 170 respectively, the boundary layer thickness δ = (A-A c ) / 2W, W is the width of the isolator, Re δ is the Reynolds number of the boundary layer thickness;

[0053] The pressure rise in the transition zone is assumed to be a discontinuity produced by a single oblique shock, and the oblique shock angle β' can be obtained from the pressure relationship before and after the oblique shock, which is:

[0054]

[0055] where π1 is the pressure ratio of the pre-combustion zone to the combustor inlet;

[0056] Since the combustion in the pre-combustion zone is mainly determined by the Mach number at the combustor inlet, the pressure ratio thereof can be fitted using the following empirical formula, which is:

[0057]

[0058] where a, b, c, d, k1, k2, e are constant coefficients, Ma t is the linear constant, Ma t1 is the turning constant 1, Ma t2 is the turning constant 2;

[0059] The length of the transition zone L is:

[0060] L = 2H / tan(β' + α)·cosα

[0061] where H is the half height of the combustor, and α is the expansion angle of the combustor;

[0062] Since the transition zone is the position of sudden pressure rise, according to the position of L0+L relative to the fuel injection hole Jet1, the combustion mode of the precombustion zone can be determined. If the position of the fuel injection hole Jet1 to the length of the combustion chamber inlet is L J1 , then: when L J1 <L0, the precombustion zone is in the mode of sustained combustion; when L0≤L J1 ≤L0+L, the precombustion zone is in the mode of transition; when L J1 >L0+L, the precombustion zone is in the mode of weak subsonic combustion. Considering the hysteresis effect existing in the hypersonic vehicle, the critical Mach number of the acceleration process can be calculated by the condition of L J1 =L0+L.

[0063] According to the critical Mach number of the combustion mode conversion of the precombustion zone, two sets of fuel injection schemes can be designed. When the Mach number of the combustion chamber inlet is lower than the critical Mach number, the precombustion zone adopts low fuel equivalence ratio for combustion, and the combustion mode is maintained in the weak subsonic combustion mode; the main combustion zone is combusted at a higher equivalence ratio, and the combustion mode is maintained in the strong subsonic combustion mode; when the Mach number of the combustion chamber inlet is higher than the critical Mach number, the fuel equivalence ratio of the main combustion zone can be appropriately increased, and the combustion mode is divided into the sustained combustion mode and the strong subsonic combustion mode. Among them, regarding the setting of the equivalence ratio of the precombustion zone and the main combustion zone, the energy and mass conservation equation sets can be established respectively to determine, specifically:

[0064] The energy and mass conservation equation sets are established in the precombustion zone as follows:

[0065]

[0066] wherein, is the mass flow rate of the combustion chamber inlet, k1 is the air-fuel ratio of the precombustion zone, P 1c is the pressure of the precombustion zone core flow, R is the gas constant, T 1c is the temperature of the precombustion zone core flow, T wall is the near-wall temperature of the precombustion zone, h1 is the height of the precombustion zone, h 1c is the height of the precombustion zone core flow, Ma 1c is the Mach number of the precombustion zone core flow, C p is the isobaric heat capacity, is the precombustion zone mass average temperature except the cavity, T is the temperature of the isolation section, ρ1 is the density of the precombustion zone, V1 is the volume of the precombustion zone, V cavity is the cavity volume, T Max is the maximum temperature in the precombustion zone cavity, η is the combustion efficiency, E fuel is the fuel heat value, ER1 is the equivalence ratio of the precombustion zone, k0 is the equivalence fuel-air ratio, L1 is the length of the precombustion zone, u1 is the average velocity of the jet wake gas of the precombustion zone;

[0067] By using the energy and mass conservation equations of the pre-combustion zone, the length L1 of the pre-combustion zone under the given equivalence ratio is solved. During the process of controlling the increase of the equivalence ratio of the pre-combustion zone, when the length L1 of the pre-combustion zone covers the cavity of the pre-combustion zone and the position 1 / 2 of the main combustion zone, it is considered that the equivalence ratio ER1 of the current pre-combustion zone has reached the required value.

[0068] The energy and mass conservation equations established in the main combustion zone are as follows:

[0069]

[0070] Among them, k2 is the air-fuel ratio of the main combustion zone, P 2c The pressure of the core flow in the main combustion zone, T 2c Temperature of the core flow in the main combustion zone, A 2c Main combustion zone core area, Ma 2c Mach number of the core flow in the main combustion zone, P2 is the pressure of the main combustion zone, T2 is the temperature of the main combustion zone, h2 is the height of the main combustion zone, ρ2 is the density of the main combustion zone, V2 is the volume of the main combustion zone, η1 is the combustion efficiency of the pre-combustion zone, η2 is the combustion efficiency of the main combustion zone, ER2 is the equivalence ratio of the main combustion zone, and L2 is the length of the main combustion zone.

[0071] By solving the energy and mass conservation equations for the main combustion zone, h1 and p1 are obtained under a given equivalence ratio. When h2 > H + x2·tanα, thermal congestion is reached, which should be avoided. That is, during the process of increasing the equivalence ratio of the main combustion zone, h2 ≤ H + x2·tanα should be maintained, where H is the half-height of the combustion chamber, x2 is the horizontal distance from the midpoint of the main combustion zone cavity to the combustion chamber expansion angle, and α is the combustion chamber expansion angle.

[0072] by Figure 1 Taking the series-connected concave-cavity combustor shown as an example, at higher and lower Mach numbers (i.e., above the critical Mach number for combustion mode transition in the pre-combustion zone and below the critical Mach number for combustion mode transition in the pre-combustion zone), the combustor should operate in the following two modes: Figure 3 The weak sub-fuel combustion-strong sub-fuel combustion mode shown and Figure 4The super-combustion and strong sub-combustion mode is shown. By the design of the combustion chamber in series, the pre-combustion zone plays the role of decelerating the gas and increasing the separation zone, thereby facilitating the full combustion of the fuel in the main combustion zone. The equivalence ratio of the fuel in the pre-combustion zone is relatively small, while the equivalence ratio of the fuel in the main combustion zone is relatively large, and the thrust of the combustion chamber is mainly generated by the main combustion zone. During the entire flight acceleration process, the combustion mode of the front cavity combustion chamber changes, while the combustion mode of the rear cavity combustion chamber remains unchanged, so that the equivalence ratio of the rear cavity combustion chamber can be adjusted to achieve a good thrust control effect. The specific injection scheme is to slowly increase the equivalence ratio of the two-stage fuel to maintain a high thrust during the low Mach number acceleration process, and to mainly increase the equivalence ratio of the main combustion zone to achieve a high thrust during the high Mach number acceleration process, since the combustion mode of the pre-combustion zone is mainly determined by the Mach number.

[0073] The specific equivalence ratio configuration scheme can be determined by the relationship between the thrust performance of the combustion chamber and the combustion equivalence ratio. First, the combustion chamber outlet parameters can be obtained, and in the main combustion zone to the outlet area, the isentropic assumption is adopted, and the gas parameters at the outlet position are calculated first, which are:

[0074]

[0075]

[0076] wherein, Ma e is the Mach number of the combustion chamber outlet, T e is the temperature of the combustion chamber outlet, p2 is the pressure of the main combustion zone, p e is the pressure of the combustion chamber outlet, is the mass flow rate of the combustion chamber outlet, A e is the cross-sectional area of the combustion chamber outlet;

[0077] Then, the thrust F generated by the combustion chamber is solved according to the thrust formula, which is:

[0078]

[0079] wherein, v e is the flow velocity of the combustion chamber outlet, P e is the pressure of the combustion chamber outlet, is the mass flow rate of the combustion chamber inlet, v in is the flow velocity of the combustion chamber inlet, P in is the pressure of the combustion chamber inlet, A in is the cross-sectional area of the combustion chamber inlet;

[0080] Finally, the fuel injection scheme of the main combustion zone can be solved according to the thrust requirement, and the process is a conventional means in the art, so this embodiment will not be described in detail.

[0081] The above merely describes the preferred embodiments of the present application, and is not intended to limit the patent scope of the present application. Any equivalent structural changes made according to the content of the present application specification and drawings, or direct / indirect application in other related technical fields, are included in the patent protection scope of the present application.

Claims

1. A method of scramjet combustor thrust control during acceleration, characterized in that, The gradually expanding series cavity combustion chamber is provided with fuel injection holes upstream of two cavities, wherein the first cavity is a pre-combustion zone and the second cavity is a main combustion zone; The thrust control method is: When the Mach number of the combustion chamber inlet is lower than the critical Mach number of the pre-combustion zone combustion mode conversion, the equivalence ratio of the pre-combustion zone is controlled to keep the combustion mode in a weak subsonic mode, and the equivalence ratio of the main combustion zone is controlled to keep the combustion mode in a strong subsonic mode; When the Mach number of the combustion chamber inlet is higher than the critical Mach number of the pre-combustion zone combustion mode conversion, the equivalence ratio of the pre-combustion zone is controlled to keep the combustion mode in a super-sonic mode, and the equivalence ratio of the main combustion zone is controlled to keep the combustion mode in a strong subsonic mode.

2. The scramjet acceleration process combustor thrust control method of claim 1, wherein, The process of controlling the equivalence ratio of the pre-combustion zone is: The energy and mass conservation equation groups are established in the pre-combustion zone, which are: wherein, is the mass flow rate of the combustion chamber inlet, k 1is the air-fuel ratio of the pre-chamber, P 1c is the pressure of the pre-chamber core flow, R is the gas constant, T 1c is the temperature of the pre-chamber core flow, T wall is the near-wall temperature of the pre-chamber, h 1is the pre-chamber height, W is the width of the isolation section, γ is the specific heat ratio, h 1c is the pre-chamber core flow height, Ma 1c is the Mach number of the pre-chamber core flow, C p is the constant pressure heat capacity, is the pre-chamber mass average temperature excluding the cavity, T is the temperature of the isolation section, ρ 1is the pre-chamber density, V 1is the volume of the pre-chamber, V cavity is the cavity volume, T Max is the maximum temperature in the pre-chamber cavity, η 1is the pre-chamber combustion efficiency, E fuel is the fuel heat value, ER 1is the equivalence ratio of the pre-chamber, k 0is the equivalence fuel-air ratio, L 1is the pre-chamber length, u 1is the pre-chamber jet wake gas average velocity; Solve the energy and mass conservation equations of the precombustion zone to obtain the length of the precombustion zone under the condition of a given equivalence ratio of the precombustion zone L 1, control the process of increasing the equivalence ratio of the precombustion zone, when the length of the precombustion zone L 1 covers the precombustion zone cavity and its position 1 / 2 to the main combustion zone, that is, the equivalence ratio of the current precombustion zone ER 1 reaches the required value.

3. The scramjet acceleration process combustor thrust control method of claim 1, wherein, In the process of controlling the equivalence ratio of the main combustion zone to keep the combustion mode in a strong subsonic mode, the equivalence ratio of the main combustion zone is prevented from being too high to cause thermal choking and prevent the thrust performance from being reduced.

4. The scramjet acceleration process combustor thrust control method of claim 1, wherein, The process of controlling the equivalence ratio of the main combustion zone is: The energy and mass conservation equation groups are established in the main combustion zone, which are: wherein, is the mass flow rate of the combustion chamber inlet, k 1 is the air-fuel ratio of the pre-chamber, k 2 is the air-fuel ratio of the main chamber, P 2c is the pressure of the main chamber core flow, R is the gas constant, T 2c is the temperature of the main chamber core flow, A 2c is the cross-sectional area of the main chamber core flow, Ma 2c is the Mach number of the main chamber core flow, γ is the specific heat ratio, P 2 is the pressure of the main chamber, T 2 is the temperature of the main chamber, W is the width of the isolation section, h 2 is the height of the main chamber, C p is the isobaric heat capacity, T is the temperature of the isolation section, ρ 2 is the density of the main chamber, V 2 is the volume of the main chamber, η 1 is the combustion efficiency of the pre-chamber, ER 1 is the equivalence ratio of the pre-chamber, η 2 is the combustion efficiency of the main chamber, ER 2 is the equivalence ratio of the main chamber, k 0 is the equivalence fuel-air ratio, E fuel is the fuel heat value, L 2 is the length of the main chamber, u 2 is the average velocity of the jet wake gas of the main chamber. solving the energy and mass conservation equations of the main combustion zone above, the equivalence ratio of the given main combustion zone is obtained h 1and p 1, the process of controlling the equivalence ratio of the main combustion zone to increase is maintained wherein, H is the half height of the combustion chamber, x 2is the horizontal distance from the midpoint of the main combustion zone cavity to the expansion corner of the combustion chamber, α is the expansion angle of the combustion chamber.

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