A method for designing a ramjet combustion chamber length

Abstract

The present application relates to a kind of ramjet combustion chamber length design method, belong to ramjet technical field.The present application includes the following steps: the flight height and flight pressure of aircraft design operating condition are given;According to the compression efficiency of inlet, design temperature rise ratio, the inlet airflow parameter is obtained in combination with isentropic relationship formula;Select fuel, the ignition delay time is obtained by using Arrhenius law to calculate chemical reaction process;The outlet airflow parameter is obtained by three conservation equations using constant pressure combustion assumption;The average speed of combustion chamber is obtained according to the mathematical average of combustion chamber inlet and outlet velocity;The combustion chamber length is obtained according to the average speed of combustion chamber and fuel ignition delay time.The present application couples the chemical power of combustion process with airflow velocity, realizes the rapid design of ramjet combustion chamber length, according to engine operating condition, combustion product can be calculated, and also according to the actual length of engine, the combustion efficiency can be inferred.

Description

Technical Field

[0001] This invention relates to a method for designing the length of a ramjet engine combustion chamber, belonging to the field of ramjet engine technology. Background Technology

[0002] Ramjet engines are ideal power sources for supersonic flight, utilizing the forward motion of the aircraft to compress air. Based on whether the airflow entering the combustion chamber is supersonic or subsonic, they can be further divided into subsonic ramjet engines and scramjet engines. The main difference lies in the Mach number; subsonic ramjet engines typically have a Mach number of around 3, while scramjet engines generally have a Mach number above 6. This difference in Mach number leads to differences in the flow field characteristics within the two types of ramjet engines, especially in the combustion organization within the combustion chamber. In subsonic ramjet engines, the airflow velocity is subsonic, and the residence time is long, allowing for mostly complete combustion; however, for some less reactive fuels, the combustion efficiency remains low. In scramjet engines, the airflow velocity is supersonic, the residence time is short, and most fuels do not burn completely.

[0003] To address the aforementioned issues, flame stabilizers or swirlers are typically used to create localized recirculation zones, extending airflow residence time and achieving thorough fuel / air mixing and combustion. However, these structures generate significant internal resistance, impacting engine performance. If increasing the combustion chamber length is used to extend airflow residence time, how is the combustion chamber length calculated? Currently, there are no reports on this in publicly available technology.

[0004] Therefore, there is an urgent need to design a method for designing the length of the combustion chamber of a ramjet engine that can solve the above-mentioned technical problems. Summary of the Invention

[0005] To address the aforementioned problems, this invention provides a method for designing the length of a ramjet engine combustion chamber. This method couples the chemical dynamics of the combustion process with the airflow velocity, enabling rapid design of the ramjet engine combustion chamber length. Based on the engine operating conditions, combustion products can be calculated, and combustion efficiency can be inferred from the actual engine length.

[0006] This invention provides a method for designing the length of a ramjet engine combustion chamber, characterized by the following steps:

[0007] Step 1: Given the flight altitude and flight pressure of the aircraft under design conditions;

[0008] Step 2: Based on the intake duct compression efficiency and design temperature rise ratio, and combined with the isentropic relationship, obtain the combustion chamber inlet airflow parameters;

[0009] Step 3: Based on the type of fuel, select typical intermediate products of the combustion process, and use Arrhenius's law to calculate the chemical reaction process to obtain the ignition delay time.

[0010] Step 4: Using the isobaric combustion assumption, the combustion chamber outlet airflow parameters are obtained through equations of mass conservation, momentum conservation, and energy conservation.

[0011] Step 5: Obtain the average velocity of the combustion chamber airflow based on the mathematical average of the inlet and outlet velocities.

[0012] Step 6: Obtain the combustion chamber length based on the average combustion chamber velocity and fuel ignition delay time.

[0013] Preferably, in step 1, the dynamic pressure of the aircraft is:

[0014]

[0015] In the formula, q0 is the flight pressure, M0 is the flight Mach number, P0 is the atmospheric pressure at the flight altitude, and γ0 is the specific heat ratio of the atmosphere at the flight altitude.

[0016] Preferably, in step 2, the combustion chamber inlet airflow parameters are:

[0017]

[0018] ρ3=P3 / (R3T3)

[0019]

[0020] In the formula, P3 is the static pressure of the combustion chamber inlet airflow, T3 is the static temperature of the combustion chamber inlet airflow, ψ is the design temperature rise ratio, and η c For intake duct compression efficiency, C p0 and C p3 U0 and U3 represent the specific heat of the incoming air at constant pressure and the specific heat of the airflow at the combustion chamber inlet, respectively. U0 and U3 represent the velocity of the incoming air and the velocity of the airflow at the combustion chamber inlet, respectively. R3 is the gas constant at the combustion chamber inlet, ρ3 is the gas density at the combustion chamber inlet, and γ3 is the specific heat ratio of the airflow at the combustion chamber inlet.

[0021] Preferably, in step 3, the selection of typical intermediate products in the combustion process is typically OH for hydrogen fuel and CH or OH for hydrocarbon fuel. The time required for their mass percentage to reach a maximum value is the ignition delay time τ.

[0022] Preferably, in step 4, the combustion chamber outlet airflow parameters are:

[0023] P4 = P3

[0024] Mass equation:

[0025] Momentum equation:

[0026] Energy equation:

[0027] In the formula, P4 is the static pressure of the gas flow at the combustion chamber outlet, T4 is the static temperature of the gas flow at the combustion chamber outlet, ρ4 is the gas density at the combustion chamber outlet, U4 represents the gas velocity at the combustion chamber outlet, and h k The enthalpy of formation of component k. The mass flow rate of the fuel. A3 represents the formation rate of component k, and A4 represents the cross-sectional area of ​​the combustion chamber inlet and outlet. In a subsonic ramjet engine, A4 / A3 is 1.0, and in a scramjet engine, it is 1.6.

[0028] Preferably, in step 5, the average velocity of the combustion chamber airflow for:

[0029]

[0030] Preferably, in step 6, the length L of the combustion chamber is:

[0031]

[0032] This invention provides a method for designing the length of a ramjet engine combustor. Given the aircraft's design operating conditions, the inlet and outlet airflow parameters of the combustor are determined sequentially to obtain the average airflow velocity. This velocity is then combined with the ignition delay time obtained from the chemical reaction calculation process to finally determine the combustor design length. During the calculation of the chemical reaction, different fuels exhibit significant differences in their fuel / air elementary reactions due to their varying chemical compositions, resulting in different combustor design lengths. This invention provides an effective method for the early-stage combustor design of ramjet engines. Attached Figure Description

[0033] Figure 1 This is the type of engine mechanism in an embodiment of the present invention;

[0034] Figure 2 This is a flowchart illustrating the method for designing the length of the combustion chamber of a ramjet engine in an embodiment of the present invention.

[0035] Figure 3 This diagram illustrates the influence of flight pressure and flight altitude on the combustion chamber length and ignition delay time, obtained using the ramjet engine combustion chamber length design method described in this embodiment.

[0036] Figure 4 This diagram illustrates the influence of the equivalence ratio on the combustion chamber length and ignition delay time, obtained using the ramjet engine combustion chamber length design method described in this embodiment.

[0037] Figure 5 The curve shows the change in the mass percentage of the intermediate product OH in the combustion chamber over time. Detailed Implementation

[0038] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0039] This embodiment discloses a method for designing the combustion chamber length of a ramjet engine, which can be used to obtain the combustion chamber length required for complete fuel combustion. (Reference) Figure 1 The figure shows the engine configuration given in this embodiment. Multiple cross-sections are divided into sections: section 0 is the incoming flow section, section 1 is the inlet section of the isolation section, section 3 is the combustion chamber inlet section, section 4 is the combustion chamber outlet section, section 9 is the outlet section of the constrained expansion section of the tail nozzle, and section 10 is the outlet section of the free expansion section of the tail nozzle. In this embodiment, the intake compression efficiency of the ramjet engine is taken as 0.9, and the temperature rise ratio is taken as 6.5.

[0040] refer to Figure 2 The method for determining the performance of a ramjet engine in this embodiment specifically includes the following steps:

[0041] Step 1: Given the design flight altitude and Mach number, calculate the flight pressure, specifically as follows:

[0042]

[0043] In the formula, q0 is the flight pressure, M0 is the flight Mach number, P0 is the atmospheric pressure at the flight altitude, and γ0 is the specific heat ratio of the atmosphere at the flight altitude.

[0044] Step 2: Based on the intake duct compression efficiency and temperature rise ratio, and combined with the isentropic relationship, the combustion chamber inlet airflow parameters are obtained, specifically:

[0045]

[0046] ρ3=P3 / (R3T3)

[0047]

[0048] In the formula, P3 is the static pressure of the combustion chamber inlet airflow, T3 is the static temperature of the combustion chamber inlet airflow, ψ is the design temperature rise ratio, and η c For intake duct compression efficiency, C p0 and C p3U0 and U3 represent the specific heat of the incoming air at constant pressure and the specific heat of the airflow at the combustion chamber inlet, respectively. U0 and U3 represent the velocity of the incoming air and the velocity of the airflow at the combustion chamber inlet, respectively. R3 is the gas constant at the combustion chamber inlet, ρ3 is the gas density at the combustion chamber inlet, and γ3 is the specific heat ratio of the airflow at the combustion chamber inlet.

[0049] Step 3: Select hydrogen as fuel and OH as a typical intermediate product. Use Arrhenius law to calculate the chemical reaction process. Set the static pressure and static temperature of the gas flow at the combustion chamber inlet as the initial pressure and temperature of the chemical reaction. The elementary reaction parameters are shown in Table 1.

[0050] Table 1: Elementary Reaction Parameters

[0051]

[0052]

[0053] Step 4: Using the isobaric combustion assumption, the combustion chamber outlet airflow parameters are obtained through the mass, momentum, and energy conservation equations, specifically:

[0054] P4 = P3

[0055]

[0056] In the formula, P4 is the static pressure of the gas flow at the combustion chamber outlet, T4 is the static temperature of the gas flow at the combustion chamber outlet, ρ4 is the gas density at the combustion chamber outlet, U4 represents the gas velocity at the combustion chamber outlet, and h k The enthalpy of formation of component k. The mass flow rate of the fuel. A represents the formation rate of component k. A3 and A4 represent the cross-sectional areas of the combustion chamber inlet and outlet, respectively, with A4 / A3 being 1.6.

[0057] Step 5, obtain the average airflow velocity in the combustion chamber, specifically:

[0058]

[0059] Step 6, obtain the design length of the combustion chamber:

[0060]

[0061] refer to Figure 3 The dashed line represents the ignition delay time, and the solid line represents the combustion chamber size. Taking H = 30 km as an example, the ignition delay time remains constant as the dynamic pressure increases. However, the combustion chamber length gradually increases. Taking q = 50 kPa as an example, the higher the flight altitude, the longer the ignition delay time and the larger the combustion chamber length.

[0062] refer to Figure 4The ignition delay time and combustion chamber length gradually increase and then decrease with the increase of the equivalence ratio. When the equivalence ratio is equal to 0.9, the ignition delay reaches its maximum value of 0.0287ms, and the combustion chamber length also reaches its maximum value of 54.68mm.

[0063] refer to Figure 5 The combustion chamber temperature rises sharply over time, eventually stabilizing at around 2500K. The OH mass percentage initially increases sharply, reaches a maximum, then gradually decreases and eventually stabilizes. Based on this, the calculated ignition delay time is 2.4 x 10⁻⁶. - 2 The oxygen mass percentage decreased rapidly from an initial 0.23 to 0.11, and then remained constant.

[0064] The above description is merely a preferred embodiment of the present invention and does not limit the patent scope of the present invention. Any equivalent structural transformations made using the contents of the present invention's specification and drawings under the inventive concept of the present invention, or direct / indirect applications in other related technical fields, are included within the patent protection scope of the present invention.

Claims

1. A method of designing the length of a ramjet engine combustion chamber, characterized in that, The method comprises the following steps: Step 1, the flight height and flight pressure of a given aircraft design working condition are given; Step 2, the inlet compression efficiency, the design temperature rise ratio and the isentropic relation are combined to obtain the combustion chamber inlet airflow parameters; Step 3, the typical intermediate product of the combustion process is selected according to the fuel type, the Arrhenius law is used to calculate the chemical reaction process, and the ignition delay time is obtained; Step 4, the isobaric combustion assumption is adopted, and the mass conservation, momentum conservation and energy conservation equations are used to obtain the combustion chamber outlet airflow parameters; Step 5, the mathematical average of the combustion chamber inlet and outlet velocities is used to obtain the combustion chamber airflow average velocity; Step 6, the combustion chamber average velocity and the fuel ignition delay time are used to obtain the combustion chamber length; The combustion chamber inlet airflow parameters are as follows: ； wherein P 3 is the static pressure of the airflow at the inlet of the combustion chamber, T 3 is the static temperature of the airflow at the inlet of the combustion chamber, ψ is the design temperature rise ratio, η c is the inlet duct compression efficiency, C p0 and C p3 represents the specific heat at constant pressure of the incoming air, the specific heat at constant pressure of the airflow at the inlet of the combustion chamber, U 0 and U 3 represents the velocity of the incoming air, the velocity of the airflow at the inlet of the combustion chamber, R 3 is the gas constant at the inlet of the combustion chamber, ρ 3 is the gas density at the inlet of the combustion chamber, γ 3 is the ratio of specific heats of the airflow at the inlet of the combustion chamber, is the atmospheric pressure at the flight altitude, is the atmospheric ratio of specific heats at the flight altitude; The time required for the mass percentage of the selected combustion intermediate products—OH for hydrogen fuel and CH or OH for hydrocarbon fuels—to reach its maximum value is defined as the ignition delay time. τ .

2. The method of claim 1, wherein In step 4, the combustion chamber outlet airflow parameters are as follows: ； Mass equation: ; Momentum equation: ; Energy equation: ; where P 4 is the static pressure of the combustion chamber exit flow, T 4 is the static temperature of the combustion chamber exit flow, ρ 4 is the density of the combustion chamber exit flow, U 4 represents the velocity of the combustion chamber exit flow, h k representing the components k of the enthalpy of formation, is the mass flow of the fuel, representing the components k of the rate of formation, A 3 and A 4 represent the cross-sectional areas at the inlet and outlet of the combustion chamber, in a ramjet, A 4 / A 3 take 1.0, in a scramjet, take 1.

6.

3. The method of claim 2, wherein: In step 5, the average velocity of the combustion chamber gas flow is: V = 0.5 * 0.5 * 0.5 。 4. The method of claim 3, wherein: In step 6, the combustion chamber length L is: 。

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