Intake, engine and exhaust matching design method for subsonic combustion ramjet engine

Abstract

The invention belongs to the field of engine inlet-engine-exhaust matching design, and particularly relates to a subsonic combustion ramjet engine inlet-engine-exhaust matching design method. The method comprises the following steps: for an initially designed air inlet channel, acquiring air inlet channel parameters; checking whether the air inlet flow meets the use requirement of the engine or not, and determining the working condition points which do not meet the requirement; the air inlet channel total pressure recovery coefficient and the flow coefficient corresponding to the working condition point are subjected to pulling deviation, and the characteristic line slope of each additional working condition is given; forming a plurality of calculation points in a designable characteristic line slope range of the air inlet channel, and determining an air inlet channel flow coefficient calculation value of each calculation point under each air inlet channel total pressure recovery coefficient deviation value; interpolating corresponding engine performance data under the additional working condition so as to construct a function relationship between the engine performance data and the air inlet total pressure recovery coefficient bias value; and according to the thrust requirement, the scheme of the air inlet channel is redesigned. According to the invention, the ramjet engine inlet-engine-exhaust matching optimization can be realized, and the use performance of the ramjet engine is improved.

Description

Technical Field

[0001] This application belongs to the field of engine intake and exhaust matching design, specifically relating to a method for matching the intake and exhaust of a sub-gas ramjet engine. Background Technology

[0002] As an air-breathing high-speed power plant, the subsonic ramjet engine has higher thrust and specific impulse than turbine engines and other power forms in the flight speed range of Mach 2 to 4. Moreover, it has a simple structure and no rotating parts, making it an ideal power choice for cruise aircraft in this speed range.

[0003] The subsonic ramjet engine consists of three parts: an air intake, a combustion chamber, and a tail nozzle. The air intake mainly decelerates and pressurizes the incoming air to achieve the airflow conditions required for the combustion chamber to operate. The combustion chamber mixes the incoming air from the air intake with the fuel carried by the engine for combustion, realizing the conversion of the fuel's chemical energy into thermal energy to heat the incoming air. The nozzle depressurizes and accelerates the high-enthalpy airflow in the combustion chamber, realizing the conversion of thermal energy into kinetic energy, thereby generating thrust and propelling the aircraft forward.

[0004] Traditional subsonic ramjet engines typically employ an integrated design of the intake manifold, combustion chamber, and exhaust nozzle, i.e., an integrated intake-engine-exhaust design, which is completed uniformly by the power unit. Its operating point is relatively fixed, the intake-engine-exhaust system lacks geometric adjustment mechanisms, the internal flow channel geometry is fixed, there are few design variables, the design objective is singular, and the performance requirements are also low. Therefore, the intake-engine-exhaust matching margin is large, and the design is more direct.

[0005] However, the design of wide-speed-range subsonic ramjet engines for aircraft faces entirely different challenges. First, due to the need for integrated flight and engine design, the aerodynamic and structural design of the exhaust system in the inlet is usually undertaken by the overall aircraft design unit, while the combustion chamber is the responsibility of the power unit. This results in data exchange and iteration between multiple units during the design process. Second, to meet the performance requirements of different flight states (such as climb and cruise) across a wide speed range, the inlet and nozzle of such engines are often equipped with geometric adjustment mechanisms (such as adjustable throats), resulting in a complex coupling relationship and a large degree of matching freedom between the intake airflow, combustion chamber inlet conditions, and engine exhaust gas coefficient.

[0006] Therefore, in the design of wide-speed-range subsonic ramjet engines for aircraft, how to establish a systematic and efficient matching process and method under the constraints that the intake, exhaust and combustion chamber are designed separately by different units and need to be optimized at multiple operating points, in order to collaboratively determine key parameters and ultimately achieve optimal engine performance in a wide speed range, especially at non-design points, has become an urgent technical problem to be solved. Summary of the Invention

[0007] To address the aforementioned issues, this application provides a method for matching the inlet and outlet of a subsonic ramjet engine. This method is geared towards the integrated design process of large-size subsonic ramjet engines by overall units in the aviation field. By meeting the integrated matching requirements of the inlet and outlet of the subsonic ramjet engine, it forms an inlet and outlet matching process and method under different constraints in a wide speed range, providing a means to optimize the thrust performance of wide-speed-range aircraft.

[0008] The inlet-outlet matching design method for a sub-gas ramjet engine provided in this application mainly includes:

[0009] Step S1: For the initial design of the air intake, obtain the air intake parameters, which include at least the height, Mach number, angle of attack, air intake flow coefficient, air intake total pressure recovery coefficient, air intake pressure ratio, air intake velocity distortion, and thrust coefficient.

[0010] Step S2: Input the intake port parameters into the engine performance database, and calculate whether the intake port flow rate based on the maximum throat area and minimum residual gas coefficient meets the engine's operating requirements. If the intake port flow rate does not meet the requirements, determine the operating point that does not meet the requirements. The operating point consists of altitude, Mach number and angle of attack.

[0011] Step S3: For each operating point that does not meet the requirements, the total pressure recovery coefficient of the intake manifold corresponding to that operating point is adjusted to obtain multiple adjustment values ​​for the total pressure recovery coefficient of the intake manifold. At the same time, the flow coefficient of the intake manifold at that operating point is adjusted to obtain multiple adjustment values ​​for the flow coefficient of the intake manifold. The multiple adjustment values ​​for the total pressure recovery coefficient of the intake manifold and the multiple adjustment values ​​for the flow coefficient of the intake manifold are combined to form additional operating conditions, and the characteristic line slope and engine performance data of each additional operating condition are given. The characteristic line slope refers to the ratio of the adjustment value for the total pressure recovery coefficient of the intake manifold to the adjustment value for the flow coefficient of the intake manifold. The engine performance data includes engine thrust and engine throat diameter.

[0012] Step S4: Form multiple calculation points within the slope range of the designable characteristic line of the intake duct, and determine the calculated value of the intake duct flow coefficient for each calculation point under the bias value of the total pressure recovery coefficient of each intake duct;

[0013] Step S5: For each intake manifold total pressure recovery coefficient pull value, interpolate the engine performance data corresponding to the calculated intake manifold flow coefficient value in the additional operating condition, thereby constructing a functional relationship between the engine performance data and the intake manifold total pressure recovery coefficient pull value.

[0014] Step S6: According to the thrust requirements, redesign the air intake scheme based on the aforementioned functional relationship. Based on the newly designed air intake, return to step S1 for iterative calculation.

[0015] Preferably, in step S1, the inlet flow coefficient is the ratio of the inlet outlet mass flow rate to the inlet reference area captured mass flow rate; the inlet total pressure recovery coefficient is the ratio of the inlet outlet total pressure to the far-field incoming total pressure; the inlet pressure boost ratio is the ratio of the inlet outlet static pressure to the far-field incoming static pressure; the inlet velocity distortion is the ratio of the difference between the maximum and minimum inlet outlet velocities to the average outlet velocity; and the thrust coefficient is the ratio of the nozzle actual thrust to the ideal thrust.

[0016] Preferably, the ideal thrust Calculated using the following formula:

[0017] ;

[0018] in, This refers to the air intake flow rate of the nozzle. Specific heat ratio of gases This is the universal gas constant. This refers to the total temperature at the nozzle inlet. The nozzle thrust ratio is the nozzle inlet pressure ratio; the actual thrust of the nozzle is the sum of the nozzle stagnation inlet impulse and the integral of the axial pressure on the nozzle wall.

[0019] Preferably, in step S3, when adjusting the intake duct total pressure recovery coefficient and intake duct flow coefficient, the adjustment amount is within ±10% of the reference value.

[0020] Preferably, the functional relationship between engine performance data and the intake manifold total pressure recovery coefficient pull value is a quadratic function.

[0021] This application obtains the relationship between the thrust characteristics of a ramjet engine and the characteristic curve of the inlet through performance biasing, total pressure recovery, and flow coefficient interpolation, and provides the constraint space for inlet characteristic curve matching, which is beneficial for further performance optimization design of the inlet. This application can achieve inlet-exhaust matching optimization of ramjet engines, improving the performance of ramjet engines. The integrated inlet structure designed based on the method of this application can achieve maximum performance of the ramjet engine during the climb phase, improving the aircraft's climb capability. Attached Figure Description

[0022] Figure 1 This is a flowchart of a preferred embodiment of the combustion chamber matching design method for a ramjet engine according to this application. Detailed Implementation

[0023] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. The described embodiments are only a part of the embodiments of this application, not all of them. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments in this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0024] This application provides a method for matching the intake and exhaust systems of a subsonic ramjet engine, such as... Figure 1 As shown, it mainly includes:

[0025] Step S1: For the initial design of the air intake, obtain the air intake parameters, which include at least the height, Mach number, angle of attack, air intake flow coefficient, air intake total pressure recovery coefficient, air intake pressure ratio, air intake velocity distortion, and thrust coefficient.

[0026] Step S2: Input the intake port parameters into the engine performance database, and calculate whether the intake port flow rate based on the maximum throat area and minimum residual gas coefficient meets the engine's operating requirements. If the intake port flow rate does not meet the requirements, determine the operating point that does not meet the requirements. The operating point consists of altitude, Mach number and angle of attack.

[0027] Step S3: For each operating point that does not meet the requirements, the total pressure recovery coefficient of the intake manifold corresponding to that operating point is adjusted to obtain multiple adjustment values ​​for the total pressure recovery coefficient of the intake manifold. At the same time, the flow coefficient of the intake manifold at that operating point is adjusted to obtain multiple adjustment values ​​for the flow coefficient of the intake manifold. The multiple adjustment values ​​for the total pressure recovery coefficient of the intake manifold and the multiple adjustment values ​​for the flow coefficient of the intake manifold are combined to form additional operating conditions, and the characteristic line slope and engine performance data of each additional operating condition are given. The characteristic line slope refers to the ratio of the adjustment value for the total pressure recovery coefficient of the intake manifold to the adjustment value for the flow coefficient of the intake manifold. The engine performance data includes engine thrust and engine throat diameter.

[0028] Step S4: Form multiple calculation points within the slope range of the designable characteristic line of the intake duct, and determine the calculated value of the intake duct flow coefficient for each calculation point under the bias value of the total pressure recovery coefficient of each intake duct;

[0029] Step S5: For each intake manifold total pressure recovery coefficient pull value, interpolate the engine performance data corresponding to the calculated intake manifold flow coefficient value in the additional operating condition, thereby constructing a functional relationship between the engine performance data and the intake manifold total pressure recovery coefficient pull value.

[0030] Step S6: According to the thrust requirements, redesign the air intake scheme based on the aforementioned functional relationship. Based on the newly designed air intake, return to step S1 for iterative calculation.

[0031] In step S1, this application first obtains the initial design scheme of the air intake, which provides the air intake parameters under different operating conditions, such as the air intake flow coefficient when the altitude is 20km, the Mach number is 3, and the angle of attack is 0. The intake manifold total pressure recovery coefficient is 0.6. The value is 0.7, and the nozzle thrust coefficient is 0.94.

[0032] In some optional embodiments, in step S1, the inlet flow coefficient is the ratio of the inlet outlet mass flow rate to the inlet reference area captured mass flow rate; the inlet total pressure recovery coefficient is the ratio of the inlet outlet total pressure to the far-field incoming total pressure; the inlet pressure boost ratio is the ratio of the inlet outlet static pressure to the far-field incoming static pressure; the inlet velocity distortion is the ratio of the difference between the maximum and minimum inlet outlet velocities to the average outlet velocity; and the thrust coefficient is the ratio of the actual nozzle thrust to the ideal thrust.

[0033] In some alternative implementations, the ideal thrust Calculated using the following formula:

[0034] ;

[0035] in, This refers to the air intake flow rate of the nozzle. Specific heat ratio of gases This is the universal gas constant. This refers to the total temperature at the nozzle inlet. The nozzle thrust ratio is the nozzle inlet pressure ratio; the actual thrust of the nozzle is the sum of the nozzle stagnation inlet impulse and the integral of the axial pressure on the nozzle wall.

[0036] Subsequently, in step S2, the aforementioned intake parameters are sent to the power unit. Based on the power unit's engine performance database, the intake flow rate under the operating conditions of maximum throat area and minimum residual gas coefficient is calculated to determine whether it meets the requirements of the ramjet engine. If it meets the requirements, it indicates that the ramjet engine can operate at maximum thrust at this operating point. If the intake flow rate does not meet the requirements, the operating point (Mach number, altitude, angle of attack) needs to be selected for secondary matching.

[0037] Understandably, there are usually tens of thousands or hundreds of thousands of operating points consisting of Mach number, altitude, and angle of attack. The operating points that do not meet the requirements are usually selected from dozens or hundreds of points, depending on the matching between the intake manifold and the engine after the initial design.

[0038] Assuming the operating conditions are 21km altitude, Mach number 3.5, and angle of attack 0, the intake flow rate does not meet the engine's operating requirements. At this point, the intake flow rate coefficient... The intake manifold total pressure recovery coefficient is 0.8. It is 0.6.

[0039] Subsequently, in step S3, the intake duct total pressure recovery coefficient and intake duct flow coefficient are adjusted.

[0040] For example, five intake flow coefficient deviation values ​​are generated, namely 0.7, 0.75, 0.8, 0.85 and 0.9, and three intake total pressure recovery coefficient deviation values ​​are generated, namely 0.5, 0.6 and 0.7.

[0041] In step S3, it can be seen that the above embodiments can form a total of 15 combinations, that is, there are 15 additional operating conditions. The altitude, Mach number and angle of attack of these additional operating conditions are unchanged, while the total pressure recovery coefficient of the intake and the flow rate coefficient of the intake are combined with each other.

[0042] Step S3 further calculates the slope of the characteristic line, which is the ratio of the deflection value of the total pressure recovery coefficient of the inlet to the deflection value of the flow coefficient of the inlet. The slope angle of the characteristic line can also be used here. To represent, that is slope angle of characteristic line The range is ~ ,in, This represents the maximum value of the intake manifold total pressure recovery coefficient pull, for example, the value of 0.7 given in the previous example. The minimum value of the intake manifold total pressure recovery coefficient pull is 0.5, as mentioned before. Similarly, This represents the maximum value of the intake flow coefficient deviation. This represents the minimum deviation value of the intake flow coefficient. Simultaneously, for the aforementioned 15 additional operating conditions, the power unit will provide a ramjet thrust performance database, i.e., engine performance data corresponding to each additional operating condition. This engine performance data must include at least engine thrust and engine throat diameter.

[0043] In some alternative implementations, in step S3, when adjusting the intake duct total pressure recovery coefficient and intake duct flow coefficient, the adjustment amount is within ±10% of the reference value.

[0044] In this embodiment, the deviation of the intake duct total pressure recovery coefficient and intake duct flow coefficient within ±10% can generally cover and match various operating conditions. The deviation of the intake duct total pressure recovery coefficient is usually less than 10%. In this case, the number can be determined according to the actual number of deviation groups. As long as the number of groups is greater than 2, it will not affect the subsequent interpolation.

[0045] Subsequently, in step S4, calculation points are selected according to the range of designable characteristic line slopes of the intake duct. For example, the designable characteristic line slope is converted into a characteristic line slope angle. The range is 0.72 to 0.87. Four calculation points are selected from this range: 0.72, 0.77, 0.82, and 0.87.

[0046] For each calculation point, the essence is the slope angle of the characteristic line. According to the three intake total pressure recovery coefficient pull values ​​given in step S3, based on the formula... This allows us to calculate three corresponding intake flow coefficients, which are referred to here as the intake flow coefficient calculation values ​​to distinguish them from the previous data.

[0047] Subsequently, in step S5, for each intake manifold total pressure recovery coefficient deflection value (a total of 3, as mentioned above, 0.5, 0.6, and 0.7 respectively), there are 4 corresponding calculation points. Each calculation point corresponds to an intake manifold flow coefficient calculation value. Based on each intake manifold flow coefficient calculation value, engine performance data (thrust F or throat diameter d) can be interpolated from the engine performance database given in step S3. Thus, the resulting data table contains the intake manifold total pressure recovery coefficient deflection value, the intake manifold flow coefficient calculation value, and engine performance data. Based on this, for each characteristic curve slope or characteristic curve slope angle... Furthermore, a functional relationship between engine performance data and the deflection value of the intake manifold total pressure recovery coefficient was constructed.

[0048] In some alternative implementations, the functional relationship between engine performance data and the intake manifold total pressure recovery coefficient pull-off value is a quadratic function.

[0049] Based on the analysis of actual impacts, constructing a quadratic equation generally yields a reasonable level of accuracy, namely: or ;in, For engine thrust, This refers to the diameter of the engine throat. , , The coefficients are those of the quadratic function.

[0050] Finally, in step S6, the above function is provided to the intake system specialist. Based on the thrust requirements, the intake system design is redesigned, and the process returns to step S1. Based on the newly designed intake system, new intake system parameters are given, initiating a new round of iterations. Through multiple iterations of the ramjet engine and intake system, an intake and exhaust matching scheme for achieving thrust in a ramjet engine can ultimately be obtained at typical operating points across a wide speed range.

[0051] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for matching the intake and exhaust manifolds of a subsonic ramjet engine, characterized in that, include: Step S1: For the initial design of the air intake, obtain the air intake parameters, which include at least the height, Mach number, angle of attack, air intake flow coefficient, air intake total pressure recovery coefficient, air intake pressure ratio, air intake velocity distortion, and thrust coefficient. Step S2: Input the intake port parameters into the engine performance database, and calculate whether the intake port flow rate based on the maximum throat area and minimum residual gas coefficient meets the engine's operating requirements. If the intake port flow rate does not meet the requirements, determine the operating point that does not meet the requirements. The operating point consists of altitude, Mach number and angle of attack. Step S3: For each operating point that does not meet the requirements, the total pressure recovery coefficient of the intake manifold corresponding to that operating point is adjusted to obtain multiple adjustment values ​​for the total pressure recovery coefficient of the intake manifold. At the same time, the flow coefficient of the intake manifold at that operating point is adjusted to obtain multiple adjustment values ​​for the flow coefficient of the intake manifold. The multiple adjustment values ​​for the total pressure recovery coefficient of the intake manifold and the multiple adjustment values ​​for the flow coefficient of the intake manifold are combined to form additional operating conditions, and the characteristic line slope and engine performance data of each additional operating condition are given. The characteristic line slope refers to the ratio of the adjustment value for the total pressure recovery coefficient of the intake manifold to the adjustment value for the flow coefficient of the intake manifold. The engine performance data includes engine thrust and engine throat diameter. Step S4: Form multiple calculation points within the slope range of the designable characteristic line of the intake duct, and determine the calculated value of the intake duct flow coefficient for each calculation point under the bias value of the total pressure recovery coefficient of each intake duct; Step S5: For each intake manifold total pressure recovery coefficient pull value, interpolate the engine performance data corresponding to the calculated intake manifold flow coefficient value in the additional operating condition, thereby constructing a functional relationship between the engine performance data and the intake manifold total pressure recovery coefficient pull value. Step S6: According to the thrust requirements, redesign the air intake scheme based on the aforementioned functional relationship. Based on the newly designed air intake, return to step S1 for iterative calculation.

2. The method for matching the intake and exhaust manifolds of a subsonic ramjet engine according to claim 1, characterized in that, In step S1, the inlet flow coefficient is the ratio of the inlet outlet mass flow rate to the inlet reference area captured mass flow rate; the inlet total pressure recovery coefficient is the ratio of the inlet outlet total pressure to the far-field incoming total pressure; the inlet pressure boost ratio is the ratio of the inlet outlet static pressure to the far-field incoming static pressure; the inlet velocity distortion is the ratio of the difference between the maximum and minimum inlet outlet velocities to the average outlet velocity; and the thrust coefficient is the ratio of the nozzle actual thrust to the ideal thrust.

3. The method for matching the intake and exhaust pipes of a subsonic ramjet engine according to claim 2, characterized in that, The ideal thrust Calculated using the following formula: ； in, This refers to the air intake flow rate of the nozzle. Specific heat ratio of gases This is the universal gas constant. This refers to the total temperature at the nozzle inlet. The nozzle thrust ratio is the nozzle inlet pressure ratio; the actual thrust of the nozzle is the sum of the nozzle stagnation inlet impulse and the integral of the axial pressure on the nozzle wall.

4. The method for matching the intake and exhaust manifolds of a subsonic ramjet engine according to claim 1, characterized in that, In step S3, when adjusting the intake duct total pressure recovery coefficient and intake duct flow coefficient, the adjustment amount is within ±10% of the reference value.

5. The method for matching the intake and exhaust manifolds of a subsonic ramjet engine according to claim 1, characterized in that, The functional relationship between engine performance data and the intake manifold total pressure recovery coefficient pull value is a quadratic function.

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