Method for calculating stress of ramjet engine inlet and exhaust component

Abstract

The application belongs to the technical field of aero-engine design calculation, and is a calculation method for stress of a ramjet engine inlet and exhaust part, wherein according to different speed and height working states of the engine and ramjet engine related parameters provided by a power unit, axial thrusts of the ramjet engine inlet, combustion chamber and tail nozzle can be obtained to realize decomposition of forces of engine parts; after the axial forces of the inlet, engine combustion chamber and tail nozzle are obtained, on the one hand, the axial forces are provided to a structure professional to facilitate the structure professional to design a plane structure frame beam according to ramjet engine force transmission requirements; on the other hand, the axial forces are provided to a control and stability professional, the control and stability professional calculates a longitudinal moment of the plane caused by axial stress of the ramjet engine combustion chamber in combination with a plane gravity center position to perform control and stability design and checking, and the method finally can meet the calculation requirements of plane structure design and flight quality related to the ramjet engine.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aero-engine design calculation, and relates to a calculation method for stress of an inlet, combustion and exhaust part of a subsonic ramjet engine. BACKGROUND

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

[0003] The subsonic ramjet engine is composed of an inlet, a combustion chamber and a nozzle. The inlet mainly pressurizes and decelerates the incoming airflow to meet the airflow conditions required by the combustion chamber. The combustion chamber mixes and burns the incoming airflow with the fuel carried by the engine to convert the chemical energy of the fuel into heat energy and heat the incoming airflow. The nozzle depressurizes and accelerates the high-enthalpy airflow to convert the heat energy into kinetic energy, thereby generating thrust to propel the aircraft forward.

[0004] The conventional subsonic ramjet engine for space missiles adopts an inlet, combustion and exhaust integrated design. The engine can serve as the barrel of the missile, which is simple in structure design and low in structural mass coefficient. However, if the subsonic ramjet engine is used as the power of an aircraft, since the aircraft structure process usually adopts a frame beam type, the barrel of the subsonic ramjet engine only serves as a part of the middle and rear fuselage of the aircraft. The combustion chamber of the engine is usually fixed by a mounting joint, and the inlet and exhaust are fixed by the frame beam. This structural layout form requires a certain prediction and evaluation capability for the stress and force transmission of each part of the subsonic ramjet engine, which is used as an input parameter for the aircraft structure design. In addition, since the thrust axis of the engine and the center of gravity of the aircraft generally do not overlap, the influence of the engine inlet, combustion and exhaust parts on the torque of the aircraft needs to be considered. SUMMARY

[0005] To solve the above problems, the application provides a calculation method for stress of an inlet, combustion and exhaust part of a subsonic ramjet engine, which comprises the following steps:

[0006] Step S1: according to the incoming airflow static temperature t0, the incoming airflow pressure p0, the working incoming airflow Mach number Ma and the specific heat ratio k, the total temperature T of the incoming airflow into the engine is calculated * and the total pressure P of the incoming airflow into the engine * and the incoming airflow velocity v0.

[0007] Step S2: according to the provided total pressure recovery coefficient δ and the inlet outlet Mach number Ma inlet the total temperature T of the incoming airflow into the engine is calculated * and the total pressure P of the incoming airflow into the engine * , the inlet outlet static temperature t is calculatedinlet , the static pressure at the inlet of the air intake p inlet and the air flow velocity at the inlet of the air intake v inlet ;

[0008] Step S3: calculate the flow rate of the incoming flow according to the flow coefficient provided obtain the fuel injection flow rate according to the excess air coefficient of the engine obtain the Mach number at the outlet of the combustion chamber Ma com , the total temperature at the outlet of the combustion chamber the total pressure at the outlet of the combustion chamber the specific heat ratio at the outlet of the combustion chamber k com , calculate the static temperature at the outlet of the combustion chamber t com , the static pressure at the outlet of the combustion chamber p com and the air flow velocity at the outlet of the combustion chamber v com ;

[0009] Step S4: calculate the axial force of the air intake F inlet based on the flow rate of the incoming flow the air flow velocity at the outlet of the air intake v inlet , the pressure of the incoming flow p0, the velocity of the incoming flow v0, the area of the air intake A inlet ;

[0010] calculate the axial force of the combustion chamber F com based on the flow rate of the outlet of the combustion chamber the air flow velocity at the outlet of the combustion chamber v com , the static pressure at the outlet of the combustion chamber p com the area of the outlet of the combustion chamber A com ;

[0011] calculate the axial force of the nozzle F noz based on the total thrust of the combustion chamber F, the axial force of the air intake F inlet , the axial force of the combustion chamber F com .

[0012] Preferably, according to the flight altitude H of the aircraft, obtain the static temperature t0, the pressure p0 and the density p0 of the incoming flow of the atmosphere at different flight altitudes based on the atmospheric function;

[0013] The specific formula includes:

[0014]

[0015]

[0016]

[0017] p0 = 100000 * p0 / R / t0;

[0018] Wherein, R is the gas constant of air.

[0019] Preferably, the formula for calculating the specific heat ratio k is:

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

[0021] Cp corresponds to the constant-pressure specific heat capacity at the flight altitude, and the formula is:

[0022] Cp=a1+a2×t0+a3×t0^2+a4×t0^3+a5×t0^4+a6×t0^5+a7×t0^6+a8×t0^7;

[0023] Wherein, a1-a8 are coefficients according to the air property values.

[0024] Preferably, based on the specific heat ratio k, the local airflow sound speed C and the incoming flow speed V0 are obtained; the formula is:

[0025]

[0026] V0=C×Ma.

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

[0028]

[0029]

[0030] Preferably, the inlet duct outlet static temperature t inlet , the inlet duct outlet static pressure p inlet , and the inlet duct outlet airflow speed v inlet are calculated, and the formula includes:

[0031]

[0032]

[0033]

[0034] Preferably, the flow rate , the oil flow rate , and the chamber outlet Mach number Ma com are calculated, and the formula includes:

[0035]

[0036]

[0037]

[0038] By q(λ), λ is solved from the following equation;

[0039]

[0040] By λ, the Mach number at the exit of the combustion chamber is calculated;

[0041]

[0042] Preferably, the axial force F inlet of the combustion chamber F com of the nozzle F noz The calculation formulae of the axial forces F

[0043]

[0044]

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

[0046] The present application is different from the design requirements of the subsonic ramjet engine used in conventional missiles. In order to solve the new requirements of structural transmission force and flight control and stability moment caused by the use of subsonic ramjet engine in aircraft, the present application innovates the calculation method of the forces of the ramjet engine components. Based on the calculation method, according to the different speed and height working states of the engine and the related parameters of the ramjet engine provided by the power unit, the axial thrusts of the inlet, the combustion chamber and the nozzle of the ramjet engine can be obtained, and the decomposition of the forces of the engine components can be realized. After obtaining the axial forces of the inlet, the combustion chamber and the nozzle of the engine, on the one hand, the forces are provided to the structure professional, so that the structure professional can design the aircraft structure frame according to the transmission force requirements of the ramjet engine; on the other hand, the forces are provided to the control and stability professional, and the control and stability professional calculates the longitudinal moment of the aircraft caused by the axial force of the combustion chamber of the ramjet engine in combination with the position of the center of gravity of the aircraft, and performs the design and checking of the control and stability. The present method can finally meet the requirements of the related structural design of the aircraft and the calculation of the flight quality of the ramjet engine. BRIEF DESCRIPTION OF DRAWINGS

[0047] Figure 1 is a flow chart of the calculation method of the forces of the inlet, the combustion chamber and the nozzle of the subsonic ramjet engine according to a preferred embodiment of the present application. DETAILED DESCRIPTION

[0048] For the technical solutions of the present application and their advantages to be clearer, the technical solutions of the present application will be further clearly and completely described in the following with reference to the drawings. It should be understood that the specific embodiments described herein are only partial embodiments of the present application, which are only used to explain the present application, rather than limit the present application. It should be noted that, for the convenience of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design. In the case of no conflict, the embodiments in the present application and the technical features in the embodiments can be combined to obtain new embodiments.

[0049] a) calculating atmospheric parameters

[0050] 1) According to the flight height H of the aircraft, the temperature, pressure, atmospheric parameters at different heights are obtained based on the atmospheric function, and the formula is as follows:

[0051]

[0052]

[0053]

[0054] p0 = 100000 x p0 / R / t0 (4)

[0055] 2) According to the formula (1) (2) (3) (4), the temperature t0, pressure p0 and density p0 in the atmosphere are obtained, and the temperature t0, pressure p0 in the atmosphere are taken as the incoming flow static temperature t0, incoming flow static pressure p0 of the aircraft. After the temperature t0 is known, the specific heat capacity Cp of the local gas can be obtained by the following formula:

[0056] Cp = a1 + a2 x t0 + a3 x t0^2 + a4 x t0^3 + a5 x t0^4 + a6 x t0^5 + a7 x t0^6 + a8 x t0^7 (5)

[0057] Wherein, since the atmospheric incoming flow static temperature is between 100-350k, the coefficients can be taken according to the air properties:

[0058] a1 = 1161.482; a2 = -2.368819; a3 = 0.01485511; a4 = -5.034909e-5; a5 = 9.9285

[0059] 69e-8; a6 = -1.111097e-10; a7 = 6.540196e14; a8 = -1.573588e-17.

[0060] 3) According to the specific heat capacity Cp corresponding to the flight height, the specific heat ratio k of the incoming air can be calculated, and the formula is 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) Based on the obtained specific heat ratio k, the local airflow velocity and the velocity of the incoming gas from the ramjet engine can be obtained, as shown in formulas (7) and (8), where C is the local sound velocity, V0 is the incoming gas velocity, and Ma is the operating incoming Mach number:

[0064]

[0065] V0=C×Ma (8)

[0066] 5) Calculate the total temperature T of the incoming flow entering the engine based on the incoming static temperature t0, incoming static pressure p0, operating incoming Mach number Ma, and specific heat ratio k. * and total pressure P * The formula is as follows:

[0067]

[0068]

[0069] b) Calculation of intake and outlet parameters

[0070] Assuming the specific heat ratio k of the entire intake remains constant, based on the total pressure recovery coefficient δ and the intake outlet Mach number Ma obtained at different speeds and altitudes provided by the intake specialist... inlet (One-dimensional mass average parameters) Calculate the static temperature t at the intake outlet. inlet Intake outlet static pressure p inlet and the airflow velocity v at the intake and exhaust ports 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 system specialist Calculate incoming flow rate The fuel injection flow rate is obtained based on the residual gas coefficient α of the engine. According to the Mach number at the combustion chamber outlet provided by the engine com Total temperature at the combustion chamber outlet Combustion chamber outlet total pressure Combustion chamber outlet specific heat ratio k com, the static temperature t com , the static pressure p com and the flow velocity v com at the exit of the combustion chamber are calculated.

[0076]

[0077]

[0078]

[0079] By q(λ), λ is calculated by the following equation.

[0080]

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

[0082]

[0083] Under the condition that the interface Mach number Ma com , the total temperature and the total pressure are known, the static temperature t com , the static pressure p com and the flow velocity v com at the exit of the engine combustion chamber are obtained according to equations (7), (8), (9) and (10).

[0084] d) Axial force calculation of each component

[0085] According to the corresponding engine operating speed height, the combustion chamber cross section related parameters, and the total thrust F of the combustion chamber, the related parameters of the inlet / outlet of the inlet / outlet of the combustion chamber, the exit / outlet of the combustion chamber are obtained by calculation, and then the axial force of the inlet F inlet , the combustion chamber F com and the tail nozzle F noz are obtained by momentum conservation, and the formula is as follows:

[0086]

[0087]

[0088] F noz = F - F inlet - F com (21).

[0089] The above merely provides the specific implementation of the present application, but the protection scope of the present application is not limited to this. Any person skilled in the art can easily think of the changes or replacements within the technical range disclosed by the present application, which should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A method for calculating the forces acting on the inlet and outlet components of a submersible ramjet engine, characterized in that, include: Step S1: Calculate the total temperature T of the incoming flow entering the engine based on the static temperature t0, the incoming pressure p0, the operating 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 provided total pressure recovery coefficient δ and inlet outlet Mach number Ma... inlet Calculate the total temperature T of the incoming flow into the engine. * and the total pressure P of the incoming flow into the engine * Calculate the static temperature t at the intake outlet. inlet Intake outlet static pressure p inlet and the airflow velocity v at the intake and exhaust ports inlet ; Step S3: Based on the provided flow coefficients Calculate incoming flow rate The fuel injection flow rate is obtained based on the residual gas coefficient α of the engine. According to the Mach number at the combustion chamber outlet provided by the engine com Total temperature at the combustion chamber outlet Combustion chamber outlet total pressure Combustion chamber outlet specific heat ratio k com Calculate the static temperature t at the combustion chamber outlet. com Combustion chamber outlet static pressure p com and combustion chamber outlet airflow velocity v com ; Step S4: Based on incoming flow Inlet outlet airflow velocity v inlet Incoming flow pressure p0, incoming flow velocity v0, inlet duct outlet area A inlet The axial force F of the intake duct is calculated. inlet ; Based on combustion chamber outlet flow Combustion chamber outlet airflow velocity v com Combustion chamber outlet static pressure p com Combustion chamber outlet A com The axial force F in the combustion chamber was calculated. com ; Based on the total thrust F of the combustion chamber and the axial force F of the intake duct inlet axial force F in the combustion chamber com The axial force F of the tail nozzle was calculated. noz .

2. The method for calculating the forces on the inlet and outlet components of a ramjet engine as described in claim 1, characterized in that, Based on the aircraft's flight altitude H and atmospheric functions, the incoming static temperature t0, incoming pressure p0, and 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 inlet and outlet components of a ramjet engine as described in claim 2, characterized in that, The formula for calculating the specific heat ratio k is: k = Cp / (Cp-R); The formula for obtaining the isobaric specific heat capacity at the flight altitude corresponding to Cp is: 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 to a8 are coefficients taken according to the physical properties of air.

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

5. The method for calculating the forces on the inlet and outlet components of a ramjet engine as described in claim 4, characterized in that, The total temperature T of the incoming flow into the engine * and the total pressure P of the incoming flow into the engine * The formula is:

6. The method for calculating the forces on the inlet and outlet components of a ramjet engine as described in claim 5, characterized in that, Intake outlet static temperature t inlet Intake outlet static pressure p inlet and the airflow velocity v at the intake and exhaust outlet inlet The calculation formula includes:

7. The method for calculating the forces on the inlet and outlet components of a ramjet engine as described in claim 6, characterized in that, Flow rate Oil flow Mach number at the combustion chamber exit com The calculation formulas include: λ can be obtained by inverse calculation using the following formula: The Mach number at the combustion chamber exit is calculated using λ.

8. The method for calculating the forces on the inlet and outlet components of a ramjet engine as described in claim 7, characterized in that, Intake axial force F inlet axial force F in the combustion chamber com Axial force F of tail nozzle noz The calculation formulas include: F noz =F-F inlet -F com 。

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