A Turbulent Combustion Simulation Method for Scramjet Engines

Abstract

This invention discloses a method for turbulent combustion simulation of scramjet engines, comprising: fitting correlation curves of temperature and pressure ratios within the isolation section of the target scramjet engine; determining multiple sump buildup pressures based on the incoming flow and obtaining the oxidizer temperature based on the correlation curves; establishing a first flamefront database coupled with multiple pressures and temperatures, and filtering the first flamefront database to obtain a second flamefront database; selecting a benchmark database and obtaining scaling factors for each second flamefront database relative to the benchmark database; and obtaining a third flamefront database matching the local pressure through linear interpolation during CFD time-progression solving, thereby completing the turbulent combustion simulation of the scramjet engine. This invention is applied to the field of scramjet engine design and can effectively handle the pressure and temperature non-uniformity effects caused by shock train compression, improving the calculation accuracy of supersonic turbulent combustion within scramjet engines.

Description

Technical Field

[0001] This invention relates to the fields of turbulent combustion simulation and scramjet engine design technology, specifically a turbulent combustion simulation method for scramjet engines. Background Technology

[0002] In recent years, hypersonic vehicle technology has become a research and focus area for major aerospace powers worldwide. The core of hypersonic vehicle technology is scramjet engine technology. Due to limitations in experimental costs and equipment, it is difficult to obtain abundant experimental data on supersonic combustion, especially since only a few experimental devices are currently available for hypersonic combustion. Therefore, numerical calculation has become an effective means of studying the internal flow and combustion of scramjet engines. Numerical calculation can obtain details of internal combustion flow that cannot be observed in experiments and conduct quantitative studies. It is an indispensable means of revealing combustion flow mechanisms, and in some cases (such as for the study of large-scale engines), it is the only effective method available. Compared with experimental measurements, numerical calculation has advantages such as convenience, speed, cost-effectiveness, the ability to reveal physical mechanisms, and rapid design optimization. In recent years, with the rapid development of computer technology and computational fluid dynamics / computation combustion dynamics (CFD / CCD), numerical calculation of the combustion flow process of scramjet engines has gradually become an engine technology research method of equal importance to ground tests and flight tests.

[0003] A key technology in numerical simulation of scramjet engines is the simulation of turbulent combustion processes. Currently, there are several main approaches: laminar finite-rate chemical reaction; PDF-based methods; and efficient flame surface methods.

[0004] The concept behind laminar finite-rate chemical reaction models is very simple and intuitive: it directly ignores the interaction between turbulence and chemical reaction, maintaining the Arrhenius form while replacing temperature and component concentration with Reynolds averages and spatially filtered values. However, neglecting the interaction between turbulence and chemical reaction often introduces significant errors, leading to decreased accuracy and failing to capture many important phenomena and processes. Scramjet engines typically exhibit a highly turbulent environment, thus laminar finite-rate chemical reaction models often present significant limitations.

[0005] The probability density function (PDF) method, starting from the stochastic nature of turbulent fluctuations, treats the turbulent flow field from a stochastic perspective. By introducing the probability density function (PDF), it derives, models, and solves the joint PDF transport equations for physical quantities to obtain single-point statistical information of the physical quantities in the turbulent field. The PDF method is theoretically complete; the exact PDF transport equations are directly derived from the Navier-Stokes equations without any additional assumptions. However, the PDF method has two significant drawbacks: firstly, numerical processing is relatively difficult; secondly, the computational cost is high, thus limiting its application in practical engineering.

[0006] Flame-face methods are widely used in engineering practice due to their advantages such as high computational efficiency and physical intuitiveness. However, traditional flame-face models are developed from subsonic combustion conditions, where the gas is incompressible. In contrast, the flow within a scramjet engine is highly compressible, and the pre-compression shockwaves generated by intense combustion further exacerbate the non-uniformity of pressure and temperature distribution in the flow field. These non-uniformities significantly affect the accuracy and reliability of flame-face models in scramjet engine turbulent combustion and must be fully considered. Figure 1 As shown, the airflow parameters at the engine combustion inlet are Ma1, p1, and T1. After shock train compression, the actual airflow parameters reaching the turbulent combustion zone should be Ma2, p2, and T2. Directly using the known inlet pressure p1 and temperature T1 to generate the flame front database will inevitably cause significant errors. Since combustion in a scramjet engine is approximately isobaric, the pressure near the entire turbulent combustion zone after the shock train is basically uniform, and the pressure before and after the flame front hardly changes. Therefore, the pressure non-uniformity correction of the flame front method can be performed based on the ratio of local pressure to inlet pressure. However, considering the influence of temperature non-uniformity is difficult because there is a large temperature gradient near the turbulent combustion zone, and the temperature difference at different locations on the flame front is very large. It is difficult to distinguish whether the local temperature change is caused by shock train compression or by the heat release of combustion itself. Existing methods focus on pressure correction, paying less attention to the temperature non-uniformity T in scramjet engines, and there is no effective method to handle the temperature non-uniformity effect. Summary of the Invention

[0007] To address the shortcomings of the prior art, this invention provides a turbulent combustion simulation method for scramjet engines, which can effectively handle the pressure and temperature non-uniformity effects caused by shock train compression.

[0008] To achieve the above objectives, the present invention provides a turbulent combustion simulation method for scramjet engines, comprising the following steps:

[0009] Step 1: Simulate the internal flow channel of the target scramjet engine and fit the correlation curves of temperature ratio and pressure ratio in the isolation section;

[0010] Step 2: Based on the incoming flow conditions, determine multiple build-up pressures within the combustion chamber pressure range, and obtain the corresponding build-up oxidant temperature based on the relevant curves and build-up pressures in Step 1.

[0011] Step 3: Use the fuel injection static temperature as the fuel end temperature for fuel storage construction, and establish a first flame surface database with multi-pressure and multi-temperature coupling based on the fuel storage pressure and corresponding oxidizer temperature in Step 2. Then, filter the first flame surface database to obtain a second flame surface database corresponding to each pressure and oxidizer temperature.

[0012] Step 4: Arbitrarily select a second flame surface database as the reference database, and obtain the scaling factor of each second flame surface database relative to the reference database.

[0013] Step 5: During the CFD time-progression solution process, based on each second flame front database and its corresponding scaling factor, a third flame front database matching the local pressure is obtained by linear interpolation according to the local pressure, thus completing the turbulent combustion simulation of the scramjet engine.

[0014] In one embodiment, step 1, specifically, involves fitting the correlation curves of the temperature ratio and pressure ratio within the isolation section, as follows:

[0015] Simulation calculations were performed on the isolation section of the target scramjet engine under different incoming Mach numbers and back pressures to obtain the ratio of the average inlet and outlet temperatures of the isolation section. Compared with back pressure ratio The distribution of the temperature ratio and pressure ratio within the isolation section was analyzed and fitted to obtain the correlation curves.

[0016] Among them, T out T represents the average temperature at the outlet of the isolation section. in P represents the average temperature at the inlet of the isolation section. out For the back pressure at the outlet of the isolation section, P in This is the back pressure at the entrance of the isolation section.

[0017] In one embodiment, step 2, determining the multiple build-up pressures within the combustion chamber pressure range, specifically involves:

[0018] Under the incoming flow condition, several friction sections are selected within the isolation section, and the pressure at each friction section is measured as the reservoir construction pressure, denoted as {P1, P2, ..., P...}. n}, where P1 <P2<···<P n P1 is the pressure at the inlet section of the isolation section, and n is the amount of reservoir construction pressure.

[0019] In one embodiment, in step 2, the temperature of the oxidant used for building the library is {T1,T2,…,T}. n}, where T1 is the average temperature of the inlet section of the isolation section;

[0020] The calculation process for the oxidant temperature of each reservoir is as follows:

[0021]

[0022] Where i = 2 to n.

[0023] In one embodiment, step 3, specifically, involves filtering the first flame surface database to obtain a second flame surface database corresponding to each pressure and oxidizer temperature, as follows:

[0024] An ensemble average of the first flamefront database was performed using a predefined PDF to obtain a second flamefront database corresponding to each pressure and oxidizer temperature. in, That is, pressure P j Oxidizing agent temperature T j The corresponding second flame surface database, j = 1 to n, where n is the number of construction pressures or construction oxidant temperatures.

[0025] In one embodiment, in step 4, the scaling factor of each second flame surface database relative to the baseline database is obtained based on the similarity of the parameter distribution, as {α1, α2, ..., α...}. n}, where α j This is the scaling factor of the j-th second flame surface database relative to the baseline database.

[0026] In one embodiment, step 5 involves obtaining a third flame face database that matches the local pressure using linear interpolation based on the local pressure. Specifically:

[0027] Obtain local pressure P local And determine the range to which the local pressure belongs within the reservoir construction pressure, i.e., P. local ∈(P k ,P k+1 ), where k = 1 to n-1;

[0028] Linear interpolation yields a third flame front database that matches the local pressure, as follows:

[0029]

[0030] In the formula, In order to meet local pressure P local A matching third flame surface database, Used as a benchmark library.

[0031] Compared with the prior art, the present invention has the following beneficial technical effects:

[0032] It is simple and easy to implement. Compared with the multi-pressure flame surface simulation method, it can take into account the effect of pre-compression of the isolation section without increasing the amount of computation and storage, thereby further improving the calculation accuracy of supersonic turbulent combustion in scramjet engines.

[0033] The physical effects are considered realistically and completely. Compared with the multi-pressure flame surface simulation method, the temperature change caused by shock train pre-compression is further considered during the database construction, which can effectively handle the pressure and temperature non-uniformity effects caused by shock train compression. Attached Figure Description

[0034] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the structures shown in these drawings without creative effort.

[0035] Figure 1 This is a schematic diagram of the internal combustion flow in a scramjet engine in the prior art;

[0036] Figure 2 This is a flowchart of the turbulent combustion simulation method in an embodiment of the present invention;

[0037] Figure 3 The following is a schematic diagram of the simulation test results of the internal flow channel of the target scramjet engine in the embodiment of the present invention, wherein: (a) is a schematic diagram of the simulation results of the isolation section under different incoming Mach numbers and different back pressures, and (b) is a schematic diagram of the correlation curves of the temperature ratio and pressure ratio in the isolation section obtained by fitting;

[0038] Figure 4 This is a schematic diagram of the correlation curves between the temperature ratio and pressure ratio within the isolation section obtained through fitting in an embodiment of the present invention.

[0039] Figure 5 In this embodiment of the invention, Y represents the Y values ​​of a benchmark database and a second flame surface database under a certain mixed fractional variance. OH Distribution map.

[0040] The realization of the objective, functional features and advantages of the present invention will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0041] 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.

[0042] Furthermore, in this invention, descriptions involving "first," "second," etc., are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0043] Furthermore, the technical solutions of the various embodiments of the present invention can be combined with each other, but only if they are feasible for those skilled in the art. If the combination of technical solutions is contradictory or cannot be implemented, it should be considered that such combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.

[0044] This embodiment discloses a turbulent combustion simulation method for scramjet engines. This method utilizes the parameter relationship before and after the shock train to correlate the temperature and pressure changes caused by the pre-compression of the shock train in the scramjet engine. It can consider the influence of the pre-compression of the isolation section without increasing the computation and storage requirements, and effectively handle the non-uniform pressure and temperature effects caused by the shock train compression, thereby improving the calculation accuracy of supersonic turbulent combustion in scramjet engines.

[0045] refer to Figure 2 The turbulent combustion simulation method in this embodiment specifically includes the following steps 1-5.

[0046] Step 1: Simulate the internal flow channel of the target scramjet engine and fit the correlation curves of temperature ratio and pressure ratio within the isolation section. The specific implementation process is as follows:

[0047] Simulation calculations were performed on the isolation section of the target scramjet engine under different incoming Mach numbers and back pressures to obtain the ratio of the average inlet and outlet temperatures of the isolation section. Compared with back pressure ratio The distribution of the temperature ratio and pressure ratio within the isolation section was analyzed and fitted to obtain the correlation curves.

[0048] Among them, T out T represents the average temperature at the outlet of the isolation section. in P represents the average temperature at the inlet of the isolation section. outFor the back pressure at the outlet of the isolation section, P in This is the back pressure at the entrance of the isolation section.

[0049] For example Figure 3-4 Simulation results of the internal flow channel of a certain scramjet engine are presented, among which... Figure 3 (a) is a contour map of the Mach number distribution within a straight isolation section with different incoming Mach numbers. Figure 3 (b) Pressure distribution cloud map within a straight isolation section at different incoming Mach numbers; Figure 4 This is a schematic diagram illustrating the correlation curves between the temperature ratio and pressure ratio within the fitted isolation section. Based on... Figure 3 The simulation results can be fitted to obtain Figure 4 The relevant curves in the figure are The coefficient of determination R of the fitted correlation curve 2 The value of 0.9 indicates a good fit. CFD calculations verified that the flame surface data with multi-pressure and multi-temperature coupling established using this fitting relationship can account for the effects of pre-compression in the isolation section without increasing computational and storage requirements, thereby further improving the computational accuracy of supersonic turbulent combustion in scramjet engines.

[0050] Step 2: Based on the incoming flow conditions, determine multiple build-up pressures within the combustion chamber pressure range, and obtain the corresponding build-up oxidant temperature based on the relevant curves and build-up pressures from Step 1. The specific implementation process is as follows:

[0051] First, under the incoming flow condition, several friction sections are selected within the isolation section, and the pressure at each friction section is measured as the reservoir construction pressure, denoted as {P1, P2, ..., P...}. n}, where P1 <P2<···<P n P1 is the pressure at the inlet section of the isolation section, and n is the quantity of reservoir construction pressure;

[0052] Based on the database construction pressure {P1,P2,…,P} n} and related curves The temperatures of the oxidant used for building the oxidant were obtained {T1,T2,…,T}. n}, where T1 is the average temperature of the inlet section of the isolation section, and the calculation process for the temperatures of the other oxidants used in the construction is as follows:

[0053]

[0054] Where i = 2 to n.

[0055] Step 3: Using the fuel injection static temperature as the fuel end temperature for fuel storage, and establishing a first flamefront database coupled with multiple pressures and temperatures based on the storage pressure and corresponding oxidizer temperature from Step 2, then filtering the first flamefront database to obtain a second flamefront database corresponding to each pressure and oxidizer temperature. Specifically, a predefined PDF is used to perform an ensemble average on the first flamefront database to obtain the second flamefront database corresponding to each pressure and oxidizer temperature. in, That is, pressure P j Oxidizing agent temperature T j The corresponding second flame surface database, j = 1 to n, where n is the number of construction pressures or construction oxidant temperatures.

[0056] It should be noted that using a predefined PDF to perform ensemble averaging on the flame surface database is a conventional technique in this field, and therefore will not be described in detail in this embodiment.

[0057] Step 4: Arbitrarily select a second flame surface database as the reference database, and obtain the scaling factor of each second flame surface database relative to the reference database. The specific implementation process is as follows:

[0058] Based on the similarity of parameter distributions, the scaling factors {α1, α2, ..., α} of each second flame surface database relative to the baseline database are obtained. n}, where α j This is the scaling factor of the j-th second flamefront database relative to the baseline database. The OH mass fraction Y in the baseline database and the second flamefront database is used as the scaling factor. OH Taking the distribution as an example, under a certain mixed fractional variance, the Y values ​​of the benchmark database and another second flame surface database... OH Distribution as Figure 5 As shown, Z on the horizontal axis represents the mixed fraction. Figure 5 The peak values ​​of the two curves have been identified. Therefore, the scaling factor of the curve in the second flame surface database relative to the curve in the baseline database is approximately 0.0039 / 0.0027 ≈ 1.444. The calculation process for the scaling factor is similar for other components and other mixture fractions.

[0059] Step 5: During the CFD time-progression solution process, based on each second flame front database and its corresponding scaling factor, a third flame front database (mixing fraction, progress variable, mixing fraction variance) matching the local pressure is obtained by linear interpolation according to the local pressure. This completes the turbulent combustion simulation of the scramjet engine. The performance of the scramjet engine can then be accurately evaluated based on the simulation results, providing a necessary foundation and basis for a deeper understanding of the complex supersonic combustion flow field and guiding the design of the combustion chamber profile.

[0060] In this embodiment, the specific implementation process of obtaining a third flame surface database that matches the local pressure based on linear interpolation of the local pressure is as follows:

[0061] First, obtain the local pressure P. local And determine the range to which the local pressure belongs within the reservoir construction pressure, i.e., P. local ∈(P k ,P k+1 ), where k = 1 to n-1;

[0062] Linear interpolation is then used to obtain a third flame face database that matches the local pressure, as follows:

[0063]

[0064] In the formula, In order to meet local pressure P local A matching third flame surface database, This serves as the second flame surface database, acting as a benchmark library.

[0065] 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 for simulating turbulent combustion in a scramjet engine, characterized in that, Includes the following steps: Step 1: Simulate the internal flow channel of the target scramjet engine and fit the correlation curves of temperature ratio and pressure ratio in the isolation section; Step 2: Based on the incoming flow conditions, determine multiple build-up pressures within the combustion chamber pressure range, and obtain the corresponding build-up oxidant temperature based on the relevant curves and build-up pressures in Step 1. Step 3: Use the fuel injection static temperature as the fuel end temperature for fuel storage construction, and establish a first flame surface database with multi-pressure and multi-temperature coupling based on the fuel storage pressure and corresponding oxidizer temperature in Step 2. Then, filter the first flame surface database to obtain a second flame surface database corresponding to each pressure and oxidizer temperature. Step 4: Arbitrarily select a second flame surface database as the reference database, and obtain the scaling factor of each second flame surface database relative to the reference database. Step 5: During the CFD time-progression solution process, based on each second flame front database and its corresponding scaling factor, a third flame front database matching the local pressure is obtained by linear interpolation according to the local pressure, thus completing the turbulent combustion simulation of the scramjet engine.

2. The turbulent combustion simulation method for scramjet engines according to claim 1, characterized in that, In step 1, the fitting process obtains the correlation curves between the temperature ratio and pressure ratio within the isolation section, specifically as follows: Simulation calculations were performed on the isolation section of the target scramjet engine under different incoming Mach numbers and back pressures to obtain the ratio of the average inlet and outlet temperatures of the isolation section. Compared with back pressure ratio The distribution of the temperature ratio and pressure ratio within the isolation section was analyzed and fitted to obtain the correlation curves. Among them, T out T represents the average temperature at the outlet of the isolation section. in P represents the average temperature at the inlet of the isolation section. out For the back pressure at the outlet of the isolation section, P in This is the back pressure at the entrance of the isolation section.

3. The turbulent combustion simulation method for scramjet engines according to claim 2, characterized in that, In step 2, determining the multiple build-up pressures within the combustion chamber pressure range specifically involves: Under the incoming flow condition, several friction sections are selected within the isolation section, and the pressure at each friction section is measured as the reservoir construction pressure, denoted as {P1, P2, ..., P...}. n }, where P1 <P2<···<P n P1 is the pressure at the inlet section of the isolation section, and n is the amount of reservoir construction pressure.

4. The turbulent combustion simulation method for scramjet engines according to claim 3, characterized in that, In step 2, the temperature of the oxidant used for building the library is {T1,T2,…,T}. n }, where T1 is the average temperature of the inlet section of the isolation section; The calculation process for the oxidant temperature of each reservoir is as follows: Where i = 2 to n.

5. The turbulent combustion simulation method for a scramjet engine according to any one of claims 1 to 4, characterized in that, In step 3, the filtering process performed on the first flame surface database to obtain the second flame surface database corresponding to each pressure and oxidizer temperature specifically involves: An ensemble average of the first flamefront database was performed using a predefined PDF to obtain a second flamefront database corresponding to each pressure and oxidizer temperature. in, That is, pressure P j Oxidizing agent temperature T j The corresponding second flame surface database, j = 1 to n, where n is the number of construction pressures or construction oxidant temperatures.

6. The turbulent combustion simulation method for a scramjet engine according to claim 5, characterized in that, In step 4, the scaling factor of each second flame surface database relative to the baseline database is obtained based on the similarity of parameter distribution, which is {α1, α2, ..., α...}. n }, where α j This is the scaling factor of the j-th second flame surface database relative to the baseline database.

7. The turbulent combustion simulation method for a scramjet engine according to claim 6, characterized in that, In step 5, a third flame front database matching the local pressure is obtained by linear interpolation based on the local pressure, specifically as follows: Obtain local pressure P local And determine the range to which the local pressure belongs within the reservoir construction pressure, i.e., P. local ∈(P k ,P k+1 ), where k = 1 to n-1; Linear interpolation yields a third flame front database that matches the local pressure, as follows: In the formula, In order to meet local pressure P local A matching third flame surface database, Used as a benchmark library.

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