Method for calculating stress of inlet, engine and exhaust parts of subsonic combustion ramjet engine

Abstract

The invention belongs to the technical field of aero-engine design and calculation, and relates to a calculation method for stress of inlet, engine and exhaust parts of a subsonic combustion ramjet, which can obtain axial thrust of an air inlet, a combustion chamber and an exhaust nozzle of the ramjet according to working states of the engine at different speeds and heights and related parameters of the ramjet provided by a power unit. The decomposition of engine part force is realized; after the axial force of the air inlet channel, the engine combustion chamber and the exhaust nozzle is obtained, on one hand, the axial force is provided for the structure specialty, and the structure specialty can conveniently design an aircraft structure frame beam according to the ramjet force transmission requirement; and on the other hand, the aircraft longitudinal moment caused by axial stress of the combustion chamber of the ramjet engine is calculated by the operation and stability specialty in combination with the gravity center position of the aircraft, operation and stability design and check are carried out, and the method can finally meet the requirements of the aircraft for structural design and flight quality calculation related to the ramjet engine.

Description

Technical Field

[0001] The present application belongs to the field of aviation engine design and calculation technology, and is a method for calculating the forces on the engine row components of a subsonic ramjet engine. Background Art

[0002] As a type of air-breathing high-speed power device, the ramjet engine has higher thrust performance and specific impulse than other power devices in the M2-4 flight speed range. Compared with turbine engines, the ramjet engine has a simple structure and no rotating parts, making it the preferred power device for cruise aircraft in the M2-4 speed range.

[0003] The ramjet engine consists of three parts: the air inlet, the combustion chamber and the tail nozzle. The air inlet is mainly used to pressurize and decelerate the incoming air to achieve the airflow conditions required for the combustion chamber to work; the combustion chamber mixes and burns the incoming air entering from the air inlet with the engine's own fuel to convert the chemical energy of the fuel into thermal energy, thereby heating the incoming air; the nozzle reduces the pressure and speeds up the high-enthalpy airflow to achieve the conversion of thermal energy into kinetic energy, thereby generating thrust and pushing the aircraft forward.

[0004] Traditional ramjet engines for space missiles adopt an integrated design of the engine and the ramjet row, so the engine can be used as the barrel of the missile, with a simple structural design and a low structural mass coefficient. However, if the ramjet engine is used as an aircraft power, since the aircraft structural process mostly adopts frame-beam type, the ramjet engine barrel is only used as a part of the aircraft's middle and rear fuselage, and the engine combustion chamber is usually fixed by mounting nodes, and the intake and exhaust are fixed by frame beams. This structural layout requires a certain prediction and evaluation capability for the force and force transmission of each component of the ramjet engine, and it is used as an input parameter for aircraft structural design. In addition, since the engine thrust axis and the aircraft's center of gravity generally do not overlap, it is necessary to consider the influence of each component of the engine's engine and the aircraft's torque. Summary of the invention

[0005] In order to solve the above problems, the present application provides a method for calculating the forces on the engine row components of a subsonic ramjet engine, comprising:

[0006] Step S1: According to the static temperature t 0 , Incoming flow pressure p 0 , the working incoming flow Mach number Ma and specific heat ratio k, calculate the total temperature T of the incoming flow entering the engine * and the total pressure P of the incoming flow into the engine * and the incoming flow velocity v 0 ;

[0007] Step S2: Based on the total pressure recovery coefficient δ and the inlet exit Mach number Ma inlet Calculate the total temperature T of the incoming air entering the engine *and the total pressure P of the incoming flow into the engine * , calculate the static temperature at the inlet duct outlet t inlet , inlet outlet static pressure p inlet and the air velocity v at the inlet outlet inlet ;

[0008] Step S3: Based on the provided flow coefficient Calculate incoming flow The injection flow rate is obtained according to the residual gas coefficient α of the engine According to the combustion chamber outlet Mach number Ma provided by the engine com , total temperature at combustion chamber outlet Total pressure at combustion chamber outlet Specific heat ratio k at combustion chamber outlet com , calculate the static temperature t at the combustion chamber outlet com , combustion chamber outlet static pressure p com and the combustion chamber outlet air velocity v com ;

[0009] Step S4: Based on the incoming flow Air flow velocity at the inlet duct outlet v inlet , Incoming flow pressure p 0 , flow velocity v 0 、Inlet duct outlet area A inlet The calculated inlet axial force F inlet ;

[0010] Based on combustion chamber outlet flow Combustion chamber outlet air velocity v com , combustion chamber outlet static pressure p com Combustion chamber outlet A com The axial force F in the combustion chamber is calculated com ;

[0011] Based on the total thrust F of the combustion chamber and the axial force F of the inlet port inlet 、Combustion chamber axial force F com The tail nozzle axial force F is calculated noz .

[0012] Preferably, according to the flight altitude H of the aircraft, based on the atmospheric function, the static temperature t of the incoming air at different flight altitudes is obtained. 0 , Incoming flow pressure p 0 and density ρ 0 ;

[0013] The specific formulas include:

[0014]

[0015]

[0016]

[0017] ρ 0 =100000×p 0 / R / t 0 ;

[0018] Where R is the gas constant of air.

[0019] Preferably, the calculation formula of the specific heat ratio k is:

[0020] k = Cp / (Cp-R);

[0021] The constant pressure specific heat capacity at the flight altitude corresponding to Cp is obtained by the formula:

[0022] Cp=a 1 +a 2 ×t 0 +a 3 ×t 0 ^2+a 4 ×t 0 ^3+a 5 ×t 0 ^4+a 6 ×t 0 ^5+a 7 ×t 0 ^6+a 8 ×t 0 ^7;

[0023] Among them, a 1 ~a 8 These are coefficients taken according to the physical properties of air.

[0024] Preferably, the local airflow sound velocity C and the incoming flow velocity V are obtained based on the specific heat ratio k. 0 ; The calculation formula is:

[0025]

[0026] V 0 =C×Ma.

[0027] Preferably, the total temperature T of the incoming air entering the engine * and the total pressure P of the incoming flow into the engine * The formula is:

[0028]

[0029]

[0030] Preferably, the static temperature at the inlet duct outlet is inlet, inlet outlet static pressure p inlet and the velocity of the airflow at the inlet duct outlet v inlet The calculation formula includes:

[0031]

[0032]

[0033]

[0034] Preferably, the flow rate Oil flow Mach number at combustion chamber exit com The calculation formula includes:

[0035]

[0036]

[0037]

[0038] Through q(λ), λ can be inversely calculated by the following formula:

[0039]

[0040] By using λ, calculate the Mach number at the combustion chamber exit;

[0041]

[0042] Preferably, the inlet axial force F inlet 、Combustion chamber axial force F com Tail nozzle axial force F noz The calculation formulas include:

[0043]

[0044]

[0045] F noz =FF inlet -F com .

[0046] The present application is different from the design requirements for the use of traditional missile subsonic ramjet engines. In order to solve the new requirements for structural force transmission and flight stability torque when the aircraft uses subsonic ramjet engines, the present invention innovates the calculation method of ramjet engine component forces; based on this calculation method, according to the working conditions of the engine at different speeds and altitudes, and the ramjet engine related parameters provided by the power unit, the axial thrust of the ramjet engine inlet, combustion chamber and tail nozzle can be obtained, thereby realizing the decomposition of the engine component forces; after obtaining the axial forces of the inlet, engine combustion chamber and tail nozzle, on the one hand, they are provided to the structure major, so that the structure major can design the aircraft structural frame beam according to the ramjet engine force transmission requirements; on the other hand, they are provided to the stability control major, and the stability control major calculates the aircraft longitudinal torque caused by the axial force of the ramjet engine combustion chamber in combination with the aircraft center of gravity position, and performs stability control design and verification. The present method can ultimately meet the aircraft's structural design and flight quality calculation requirements related to ramjet engines. BRIEF DESCRIPTION OF THE DRAWINGS

[0047] Figure 1 It is a flow chart of a method for calculating the forces on the engine row components of a sub-combustion ramjet engine in a preferred embodiment of the present application. DETAILED DESCRIPTION

[0048] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be further described in detail in detail and in detail in conjunction with the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0049] a) Calculation of atmospheric parameters

[0050] 1) According to the aircraft's flight altitude H, based on the atmospheric function, the temperature, pressure, and atmospheric parameters at different altitudes are obtained. The formula is as follows:

[0051]

[0052]

[0053]

[0054] ρ 0 =100000×p 0 / R / t 0 (4)

[0055] 2) According to formula (1)(2)(3)(4), the temperature t in the atmosphere is obtained:0 , pressure p 0 and density ρ 0 , the atmospheric temperature t 0 , pressure p 0 As the static temperature of the incoming air entering the aircraft 0 , static pressure of incoming flow p 0 , at a known temperature t 0 After that, the constant pressure specific heat capacity Cp of the local gas can be obtained by the following formula:

[0056] Cp=a 1 +a 2 ×t 0 +a 3 ×t 0 ^2+a 4 ×t 0 ^3+a 5 ×t 0 ^4+a 6 ×t 0 ^5+a 7 ×t 0 ^6+a 8 ×t 0 ^7 (5)

[0057] Since the static temperature of the atmospheric flow is between 100-350K, the coefficient can be taken according to the physical properties of air:

[0058] a 1 =1161.482; a 2 =-2.368819; a 3 =0.01485511; a 4 =-5.034909e-5; a 5 =9.9285

[0059] 69e-8;a 6 =-1.111097e-10; a 7 =6.540196e14;a 8 =-1.573588e-17.

[0060] 3) According to the constant pressure specific heat capacity Cp at the corresponding flight altitude, the specific heat ratio k of the incoming air can be calculated as follows:

[0061] a) k = Cp / (Cp-R) (6)

[0062] Where R is the gas constant of air, which is taken as 287;

[0063] 4) According to the obtained specific heat ratio k, the local airflow speed and the speed of the ramjet engine flow gas can be obtained, as shown in formulas (7) and (8), where C is the local sound speed, V 0 is the incoming flow velocity, Ma is the working incoming flow Mach number:

[0064]

[0065] V 0 =C×Ma (8)

[0066] 5) According to the static temperature of the incoming flow t 0 , static pressure of incoming flow p 0 , the working incoming flow Mach number Ma and specific heat ratio k, calculate the total temperature T of the incoming flow entering the engine * and total pressure P * The formula is as follows:

[0067]

[0068]

[0069] b) Calculation of inlet outlet parameters

[0070] Assuming that the specific heat ratio k of the entire inlet duct remains unchanged, according to the total pressure recovery coefficient δ and the inlet outlet Mach number Ma obtained at different inlet speeds and altitudes provided by the inlet professional inlet (one-dimensional mass average parameter), calculate the static temperature t at the inlet outlet inlet , inlet outlet static pressure p inlet and the air velocity v at the inlet outlet inlet ; The formula is as follows:

[0071]

[0072]

[0073]

[0074] c) Calculation of combustion chamber outlet parameters

[0075] According to the flow coefficient provided by the intake duct professional Calculate incoming flow The injection flow rate is obtained according to the residual gas coefficient α of the engine According to the combustion chamber outlet Mach number Ma provided by the engine com , total temperature at combustion chamber outlet Total pressure at combustion chamber outlet Specific heat ratio k at combustion chamber outlet com , calculate the static temperature t at the combustion chamber outlet com , combustion chamber outlet static pressure p comand the combustion chamber outlet air velocity v com .

[0076]

[0077]

[0078]

[0079] Through q(λ), λ can be inversely calculated using the following formula.

[0080]

[0081] By using λ, the Mach number at the combustion chamber exit is calculated.

[0082]

[0083] At a known interface Mach number Ma com , total temperature Total pressure Under the condition of , according to formula (7)(8)(9)(10), the static temperature t at the engine combustion chamber outlet is obtained com , combustion chamber outlet static pressure p com and the combustion chamber outlet air velocity v com .

[0084] d) Calculation of axial forces of each component

[0085] According to the corresponding engine operating speed height, combustion chamber section parameters, and the total thrust F of the combustion chamber, the relevant parameters of the inlet duct outlet / combustion chamber inlet and combustion chamber outlet / nozzle inlet sections were obtained by calculation, and then the inlet duct F was obtained by conservation of momentum. inlet 、Combustion chamber F com and tail nozzle F noz The axial force is given by the following formula:

[0086]

[0087]

[0088] F noz =FF inlet -F com (twenty one).

[0089] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A method for calculating the forces on the engine row components of a subsonic ramjet engine, characterized in that: include: Step S1: Calculate the total temperature T of the incoming flow entering the engine according to the incoming flow static temperature t0, the incoming flow pressure p0, the working incoming flow Mach number Ma and the specific heat ratio k. * and the total pressure P of the incoming flow into the engine * and the incoming flow velocity v0; Step S2: Based on the total pressure recovery coefficient δ and the inlet exit Mach number Ma inlet Calculate the total temperature T of the incoming air entering the engine * and the total pressure P of the incoming flow into the engine * , calculate the static temperature at the inlet duct outlet t inlet , inlet outlet static pressure p inlet and the air velocity v at the inlet outlet inlet ; Step S3: Based on the provided flow coefficient Calculate incoming flow The injection flow rate is obtained according to the residual gas coefficient α of the engine According to the combustion chamber outlet Mach number Ma provided by the engine com , total temperature at combustion chamber outlet Total pressure at combustion chamber outlet Specific heat ratio k at combustion chamber outlet com , calculate the static temperature t at the combustion chamber outlet com , combustion chamber outlet static pressure p com and the combustion chamber outlet air velocity v com ; Step S4: Based on the incoming flow Air flow velocity at the inlet duct outlet v inlet , inlet pressure p0, inlet velocity v0, inlet duct outlet area A inlet The calculated inlet axial force F inlet ; Based on combustion chamber outlet flow Combustion chamber outlet air velocity v com , combustion chamber outlet static pressure p com Combustion chamber outlet A com The axial force F in the combustion chamber is calculated com ; Based on the total thrust F of the combustion chamber and the axial force F of the inlet port inlet , combustion chamber axial force F com The tail nozzle axial force F is calculated noz .

2. The method for calculating the forces on the engine row components of a subsonic ramjet engine according to claim 1, characterized in that: According to the aircraft's flight altitude H, based on the atmospheric function, the static temperature t0, the pressure p0 and the density ρ0 of the atmosphere at different flight altitudes are obtained; The specific formulas include: ρ0=100000×p0 / R / t0; Where R is the gas constant of air.

3. The method for calculating the forces on the engine row components of a subsonic ramjet engine according to claim 2, characterized in that: The calculation formula of specific heat ratio k is: k = Cp / (Cp-R); The constant pressure specific heat capacity at the flight altitude corresponding to Cp is obtained by the formula: Cp=a1+a2×t0+a3×t0^2+a4×t0^3+a5×t0^4+a6×t0^5+a7×t0^6+a8×t0^7; Among them, a1~a8 are coefficients determined according to the physical properties of air.

4. The method for calculating the forces on the engine row components of a subsonic ramjet engine according to claim 3, characterized in that: Based on the specific heat ratio k, the local airflow sound velocity C and the incoming flow velocity V0 are obtained; the calculation formula is: V0=C×Ma.

5. The method for calculating the forces on the engine row components of a subsonic ramjet engine according to claim 4, characterized in that: The total temperature of the incoming flow entering the engine is T * and the total pressure P of the incoming flow into the engine * The formula is:

6. The method for calculating the forces on the engine row components of a subsonic ramjet engine according to claim 5, characterized in that: Inlet outlet static temperature t inlet , inlet outlet static pressure p inlet and the velocity v of the airflow at the inlet duct outlet inlet The calculation formula includes:

7. The method for calculating the forces on the engine row components of a subsonic ramjet engine according to claim 6, characterized in that: Flow rate Oil flow Mach number at combustion chamber exit com The calculation formula includes: Through q(λ), λ can be inversely calculated by the following formula: By using λ, calculate the Mach number at the combustion chamber exit; 8. The method for calculating the forces on the engine row components of a subsonic ramjet engine as claimed in claim 7, characterized in that: Inlet axial force F inlet , combustion chamber axial force F com Tail nozzle axial force F noz The calculation formulas include: F noz =F-F inlet -F com 。

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