Modeling method of component-level supersonic air inlet comprehensive model based on aero-engine

Abstract

The invention discloses a component-level supersonic air inlet comprehensive model modeling method based on an aero-engine. The method comprises the steps that initialization parameters are acquired to define an initial model; in the initial model, constructing an air inlet component-level model; in the initial model, constructing an engine part-level model; the combination of the air inlet component-level model and the engine component-level model is regarded as a comprehensive dynamic model, and the comprehensive dynamic model is solved by using a cooperative work equation set; using a trust region Newton first pass algorithm to optimize the cooperative work equation set so as to update guess value parameters of the comprehensive dynamic model; according to the method, through coupling of the air inlet component-level model and the engine component-level model, accurate modeling is carried out for overflow resistance change and air mass flow dynamic adjustment during supersonic flight, and the stability of the propulsion system under different working conditions is effectively enhanced; and the performance loss or fault risk caused by mismatching of the air inlet channel and the engine is reduced.

Description

Technical Field

[0001] The invention relates to an aero-engine, in particular to a modeling method for a component-level supersonic inlet comprehensive model based on an aero-engine. Background Art

[0002] In recent years, with the expansion of the flight envelope of fixed-wing aircraft, especially the growing performance requirements under supersonic flight conditions, higher requirements have been placed on the comprehensive modeling methods of aviation propulsion systems. Although there are relatively rich studies on independent supersonic inlet modeling and engine component-level models, there is a relative lack of methodologies for the coupling characteristics of the two and the comprehensive modeling of the overall propulsion system.

[0003] Under supersonic flight conditions, the working state of the inlet is directly related to the efficiency and stability of the entire propulsion system. At this time, the distortion phenomenon at the inlet outlet, the increase in overflow resistance, and the air mass flow matching problem will significantly affect the installation performance of the engine. Traditionally, the engine component-level model focuses on the detailed description of core components such as the fan, compressor, combustion chamber, high-pressure turbine, low-pressure turbine and tail nozzle, while often ignoring the importance of the inlet component-level model. In the existing engine modeling practice, the research on the deep-level inlet-engine coupling mechanism is still limited. For example, the patent contribution with patent publication number CN116720261A discloses a method for establishing an aero-engine component-level model with the ability to simulate performance uncertainty. Although it can simulate the component-level model of performance uncertainty during production and use, this method may not accurately simulate the behavior under incomplete matching conditions when the aircraft is decelerated or other non-ideal conditions, especially when the propulsion system generates additional overflow resistance.

[0004] To this end, the present invention provides a modeling method for a component-level supersonic inlet comprehensive model based on an aircraft engine. Summary of the invention

[0005] In view of the deficiencies in the prior art, the present invention provides a component-level supersonic inlet comprehensive model modeling method based on aircraft engines, which solves the flow matching problem between the engine and the inlet that is difficult to simulate through a trust region Newton one-pass method.

[0006] To achieve the above objectives, the present invention is implemented through the following technical solutions:

[0007] The modeling method of the component-level supersonic inlet comprehensive model based on aero-engine includes:

[0008] S1, obtain initialization parameters to define the initial model;

[0009] S2. In the initial model, construct an intake duct component level model;

[0010] S3. In the initial model, construct an engine component level model;

[0011] S4, treating the combination of the intake duct component-level model and the engine component-level model as a comprehensive dynamic model, and solving the comprehensive dynamic model using a collaborative working equation group;

[0012] S5. Optimizing the collaborative working equations using the trust region Newton one-pass algorithm to update the guessed parameters of the comprehensive dynamic model;

[0013] The guessed parameters include: inlet flow coefficient, fan pressure ratio coefficient, compressor pressure ratio coefficient, high-pressure turbine inlet flow and low-pressure turbine inlet flow.

[0014] In some embodiments, the step of defining the initial model includes:

[0015] S1-1, obtaining initialization parameters, including: flight altitude H, flight Mach number Ma, fuel quantity, nozzle outlet area, fan pressure ratio coefficient, compressor pressure ratio coefficient;

[0016] S2, defining the initial state of the component-level model according to the initialization parameters;

[0017] In some of the embodiments, in the initial model, building an intake duct component level model includes:

[0018] S2-1, construct the inlet flow model;

[0019] S2-2, construct the low-speed internal flow model of the inlet;

[0020] S2-3. Construct an inlet outflow model.

[0021] In some of the embodiments, constructing an inlet flow model includes:

[0022] S2-1-1. Obtain the total temperature and flight Mach number in the flight state, and substitute them into the following formula to generate the total temperature at the inlet outlet;

[0023]

[0024] Among them, T0 is the total temperature in flight, Ma is the flight Mach number, and κ is the air adiabatic index;

[0025] S2-1-2. Obtain the total pressure and flight Mach number in the flight state, substitute them into the following formula to generate the total pressure at the inlet outlet;

[0026]

[0027] Among them, p0 is the total pressure in flight state, and p2 is the total pressure at the inlet outlet;

[0028] S2-1-3, obtaining the total temperature and total pressure at the inlet duct outlet, and generating the air density at the inlet duct outlet;

[0029] The expression for calculating the air density at the inlet duct outlet is:

[0030] Among them, ρ2 is the air density at the inlet outlet, and R is the gas constant;

[0031] S2-1-4. Obtain the outlet cross-sectional area, velocity coefficient, and flow function, and substitute them into the following formula to calculate the air mass flow rate;

[0032] The expression for calculating air mass flow is:

[0033] in, is the air mass flow rate at the inlet duct outlet, A2 is the cross-sectional area of ​​the inlet duct outlet, λ2 is the inlet duct outlet velocity coefficient, and q(λ2) is the inlet duct outlet flow coefficient.

[0034] S2-1-5. Based on the calculated total temperature at the inlet duct outlet, total pressure at the inlet duct outlet and air mass flow rate, establish a flow parameter distribution to construct the inlet duct internal flow model.

[0035] In some of the embodiments, constructing an inlet low-speed internal flow model includes:

[0036] S2-2-1. When the flight Mach number Ma is less than 0.6, the total pressure recovery coefficient is calculated using the empirical formula;

[0037] σ=0.782+0.147cos(Ma)+0.152sin(Ma);

[0038] Where σ is the total pressure recovery coefficient.

[0039] In some of the embodiments, constructing an inlet outflow model includes:

[0040] S2-3-1. Obtain the overflow resistance, venting resistance and boundary layer discharge resistance of the air inlet;

[0041] S2-3-2, inputting the overflow resistance, the deflation resistance and the boundary layer discharge resistance of the inlet duct, and calculating the total outflow resistance;

[0042] The expression for calculating the sum of the outflow resistance is: in =D sp +D bp +D bl

[0043] Among them, Dsp is the overflow resistance, which describes the overflow characteristics of the excess air in the intake duct; D bp The deflation resistance describes the resistance of the external airflow caused by the deflation characteristics; D bl The boundary layer shedding resistance describes the resistance caused by the shedding effect of the boundary layer.

[0044] S2-3-3, obtain the total external flow resistance, flight Mach number, inlet cross-sectional area, and specific heat ratio of air;

[0045] S2-3-4. Calculate the inlet drag coefficient by inputting the total outflow resistance, flight Mach number, inlet cross-sectional area, and specific heat ratio of air;

[0046] The expression for calculating the inlet resistance coefficient is:

[0047] Among them, C in is the inlet resistance coefficient, D in is the total outflow resistance, γ is the specific heat ratio of air, A c is the inlet cross-sectional area, p0 is the total pressure of the air;

[0048] S2-3-5. Obtain the drag coefficient, flight speed, air density and cross-sectional area;

[0049] S2-3-6. Calculate the total external resistance of the air inlet and tail nozzle according to the drag coefficient, flight speed, air density and cross-sectional area;

[0050] The expression for calculating the total external resistance of the air inlet and the tail nozzle is:

[0051] Among them, F inlet,drag represents the total external resistance of the air inlet and tail nozzle, ρ is the air density, v is the flight speed, A c is the cross-sectional area of ​​the inlet duct.

[0052] In some of the embodiments, in the initial model, an engine component level model is constructed, including:

[0053] S3-1, obtaining the converted flow rate at the inlet duct outlet and the converted flow rate at the engine fan, and constructing a residual equation for matching and balancing the inlet duct and engine flow rates;

[0054] The residual equation for the balance between the intake duct and the engine flow is:

[0055] Among them, m 2,c Indicates the calculated flow rate at the inlet duct outlet, m 21,c represents the calculated flow rate of the engine fan, ∈1 represents the residual between the intake duct and the engine flow rate;

[0056] S3-2, obtaining the high-pressure turbine guide valve inlet flow rate and the high-pressure turbine characteristic line to calculate the flow rate, and constructing the residual equation of the high-pressure turbine rotor inlet flow balance;

[0057] The residual equation for the high-pressure turbine rotor inlet flow balance is:

[0058] Among them, m 41,cx represents the calculated inlet flow rate of the high-pressure turbine guide, m 41,c represents the actual inlet flow of the high-pressure turbine, ∈2 represents the residual between the inlet flow of the high-pressure turbine guide valve and the flow calculated by the high-pressure turbine characteristic line;

[0059] S3-3, obtaining the flow rate of the low-pressure turbine guide vane inlet and the flow rate calculated by the high-pressure turbine characteristic line, and constructing the residual equation of the low-pressure turbine rotor inlet flow balance;

[0060] The residual equation for the low-pressure turbine rotor inlet flow balance is:

[0061] Among them, m 45,cx represents the calculated inlet flow rate of the low-pressure turbine guide vane, m 45,c represents the actual inlet flow of the low-pressure turbine, ∈3 represents the residual between the inlet flow of the low-pressure turbine guide valve and the flow calculated by the high-pressure turbine characteristic line;

[0062] S3-4, obtaining the static pressure at the inner and outer outlets of the engine, and constructing the residual equation of the static pressure balance at the inlet of the mixing chamber;

[0063] The residual equation of static pressure balance at the mixing chamber inlet is:

[0064] Among them, p 16,s represents the static pressure at the outlet of the vortex mixing chamber in the engine, p 6,s represents the static pressure of the outer vortex mixing chamber, ∈4 represents the residual between the static pressure at the engine outlet and the static pressure at the outer vortex outlet;

[0065] S3-5, obtain the flow rate of the tail nozzle and the flow characteristics of the tail nozzle to calculate the flow rate, and construct the residual equation of the tail nozzle throat flow balance;

[0066] The residual equation of the tail nozzle throat flow balance is:

[0067] Among them, m 8,x represents the calculated flow rate of the tail nozzle, m8 represents the actual flow rate of the tail nozzle, ∈5 represents the residual between the flow rate of the tail nozzle and the flow rate calculated by the internal flow characteristics of the tail nozzle;

[0068] S3-6, obtaining the power consumed by the fan and the power provided by the low-pressure turbine, and constructing a residual equation for the power balance of the low-pressure rotor;

[0069] The residual equation for the low-pressure rotor power balance is:

[0070] Among them, W f Indicates the power consumed by the fan, W lt represents the power provided by the low-pressure turbine, η represents the mechanical transmission efficiency, ∈6 represents the residual between the power consumed by the fan and the power provided by the low-pressure turbine;

[0071] S3-7, obtaining the power consumed by the fan and the power provided by the high-pressure turbine, and constructing a residual equation balanced by the power of the high-pressure rotor;

[0072] The residual equation for the high-pressure rotor power balance is:

[0073] Among them, W c represents the power consumed by the compressor, Wht represents the power provided by the high-pressure turbine, η h represents the mechanical transmission efficiency of the high-pressure rotor shaft, ∈7 represents the residual between the power consumed by the fan and the power provided by the high-pressure turbine

[0074] S3-8, combining the residual variances from S3-1 to S3-7 to obtain the collaborative engineering equations;

[0075] S3-9. Use the dynamic characteristic equation to calculate the speed change rate of the high and low pressure rotors under dynamic conditions;

[0076]

[0077] Among them, N f and N c are the speeds of the fan and compressor respectively, N lt and N ht are the speeds of the low-pressure turbine and the high-pressure turbine, respectively, and J f and J c are the rotational inertia of the fan and compressor respectively;

[0078] S3-10, combining the collaborative engineering group with the dynamic characteristic equation to form the engine component level model

[0079] In some embodiments, solving the intake duct component level model and the engine component level model using the cooperative working equations includes:

[0080] S4-1, combining the intake duct component level model and the engine component level model to construct a comprehensive dynamic model;

[0081] S4-2, selecting initialization parameters to perform iterative updates on the comprehensive dynamic model;

[0082] S4-3, when the norm of the residual matrix is ​​updated iteratively to meet the convergence threshold, the output balance of the comprehensive dynamic model is determined, otherwise the iteration continues;

[0083] ∈ n =[∈1,∈2,…,∈7] T

[0084] ||∈ n ||≤δ

[0085] Among them, ∈ n represents the residual matrix, ||∈ n || represents the norm of the residual matrix, and δ represents the convergence threshold.

[0086] In some embodiments, a trust region Newton one-pass algorithm is used to optimize the collaborative working equations to update the guessed parameters of the comprehensive dynamic model, including:

[0087] Using the Trust Region Newton One-Pass Algorithm, the objective function of the collaborative engineering equations is defined;

[0088] The trust region Newton one-pass algorithm passes through the iteration point x k An approximate objective optimization function of the objective function f(x) is constructed to solve the problem;

[0089] Among them, the approximate objective optimization function is:

[0090]

[0091] Among them, s k =x k+1 -x k ;

[0092] s k is the step size, which means from the current iteration point x k To the next iteration point x k+1 The displacement is used to optimize the update of variables; x k is the current iteration point, indicating the location of the current optimization solution, x k+1 is the next iteration point, indicating the new solution updated after this iteration; the step definition formula describes the step length between the current solution and the next solution during the iteration process, and the step length s k It is the core of variable update.

[0093] F k (s k ) is the approximate objective function, which is constructed by quadratic approximation f(x) and is used to find the optimal solution in the current trust region; f(x k) is the objective function f(x) at the current iteration point x k The function value of g k is the objective function f(x) at the iteration point x k The gradient vector at is defined as:

[0094]

[0095] The gradient vector describes the direction and rate of change of the objective function at the current iteration point; is the transpose of the gradient vector, which changes the gradient vector from a column vector to a row vector for use with the step size s k Perform inner product operation; is a linear term, indicating that the objective function is at the current point x k Along step length s k The first-order change in direction;

[0096] H k is the objective function f(x) at x k The Hessian matrix at is defined as the second-order partial derivative matrix:

[0097]

[0098] is a quadratic term used to adjust the weight of the quadratic approximation model; ||s k || is the step length s k The second norm of is defined as: It represents the length of the step; h k is the trust region radius, which indicates the upper bound of the step size in the current iteration; it is used to limit the step size s k The size of , thus ensuring that the optimization solution converges within the trust region; ||s k ||≤h k As a constraint condition, the step length s is required k The length of the trust region does not exceed the radius h k ;

[0099] In some embodiments, the trust region is defined as a given trust region radius h k Inlet flow coefficient φ in,k Neighborhood of;

[0100] The definition of the trust region is:

[0101]

[0102] Among them, Ω k represents the trust region in the kth iteration; the trust region is defined as the optimization variable x restricted to the iteration point φ in,k The area within a certain range nearby is determined by the trust region radius hk Decide;

[0103] x represents the possible variable value in the optimization process; φ in,k represents the inlet flow coefficient at the current iteration point, which is the solution of the trust region Newton one-pass algorithm in the kth iteration;

[0104] ||x-φ in,k || represents vector x and φ in,k The distance between is defined using the two-norm: Among them, x i and φ in,k are vectors x and φ respectively in,k The i-th component of .

[0105] The present invention provides a component-level supersonic inlet comprehensive model modeling method based on an aircraft engine, which has the following beneficial effects:

[0106] Through the coupling of the inlet component-level model and the engine component-level model proposed in the present invention, accurate modeling is performed for the overflow resistance change and the dynamic adjustment of the air mass flow during supersonic flight, which helps to enhance the stability and efficiency of the propulsion system under different operating conditions and reduce the performance loss or failure risk caused by the mismatch between the inlet and the engine.

[0107] In addition, by using a set of collaborative working equations and a trust-region Newton one-pass algorithm for optimization, the optimal solution for guessed parameters such as overflow resistance change and air mass flow rate can be found more quickly and accurately, further improving the overall performance of the propulsion system. BRIEF DESCRIPTION OF THE DRAWINGS

[0108] Figure 1 A schematic diagram of a process for updating guessed value parameters of a comprehensive dynamic model of the present invention;

[0109] Figure 2 A schematic diagram of condition determination for iterative updating of the present invention;

[0110] Figure 3 is a structural block diagram of the air intake duct component level model of the present invention;

[0111] Figure 4 It is a structural block diagram of the engine component level model of the present invention. DETAILED DESCRIPTION

[0112] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0113] Example 1: Please refer to Figure 1-Figure 4 The present invention provides a method for modeling a component-level supersonic inlet comprehensive model based on an aircraft engine, comprising the following step S1:

[0114] S1, obtain initialization parameters to define the initial model;

[0115] In step S1, the initial model definition step includes:

[0116] S1-1, obtaining initialization parameters, including: flight altitude H, flight Mach number Ma, fuel quantity, nozzle outlet area, fan pressure ratio coefficient, compressor pressure ratio coefficient;

[0117] S2, defining the initial state of the component-level model according to the initialization parameters;

[0118] Among them, the fuel volume represents the mass flow rate of fuel consumed by the engine per unit time, which is usually used to evaluate the engine's working efficiency and thrust level; the nozzle outlet area represents the outlet cross-sectional area of ​​the tail nozzle, which directly affects the tail nozzle flow and jet velocity, thereby affecting the thrust; the fan pressure ratio coefficient represents the ratio of the engine fan airflow to the core airflow, which is used to measure the efficiency of the fan; the compressor pressure ratio coefficient represents the ratio of the compressor outlet pressure to the inlet pressure, which is an important parameter of the engine's working state and is used to describe the compressor's boost capability.

[0119] Specifically, the initialization process directly affects the modeling accuracy. By inputting parameters such as flight altitude and Mach number to calculate the initial flow field distribution, combined with key variables such as fuel quantity, the aerodynamic and thermal states of the air inlet and engine are initially set, providing accurate input for subsequent model construction.

[0120] The modeling method further comprises step S2:

[0121] S2. In the initial model, construct an intake duct component level model;

[0122] In step S2, in the initial model, an intake duct component level model is constructed, including:

[0123] S2-1, construct the inlet flow model;

[0124] In step S2-1, an inlet flow model is constructed, including:

[0125] S2-1-1. Obtain the total temperature and flight Mach number in the flight state, and substitute them into the following formula to generate the total temperature at the inlet outlet;

[0126]

[0127] Among them, T0 is the total temperature in flight, Ma is the flight Mach number, and k is the air adiabatic index;

[0128] S2-1-2. Obtain the total pressure and flight Mach number in the flight state, substitute them into the following formula to generate the total pressure at the inlet outlet;

[0129]

[0130] Among them, p0 is the total pressure in flight state, and p2 is the total pressure at the inlet outlet;

[0131] S2-1-3, obtaining the total temperature and total pressure at the inlet duct outlet, and generating the air density at the inlet duct outlet;

[0132] The expression for calculating the air density at the inlet duct outlet is:

[0133] Among them, ρ2 is the air density at the inlet outlet, and R is the gas constant;

[0134] S2-1-4. Obtain the outlet cross-sectional area, velocity coefficient, and flow function, and substitute them into the following formula to calculate the air mass flow rate;

[0135] The expression for calculating air mass flow is:

[0136] in, is the air mass flow rate at the inlet duct outlet, A2 is the cross-sectional area of ​​the inlet duct outlet, λ2 is the inlet duct outlet velocity coefficient, and q(λ2) is the inlet duct outlet flow coefficient.

[0137] S2-1-5. Based on the calculated total temperature at the inlet duct outlet, total pressure at the inlet duct outlet and air mass flow rate, establish a flow parameter distribution to construct the inlet duct internal flow model.

[0138] Specifically, when constructing the inlet duct internal flow model, the outlet parameters are first calculated using the total temperature and total pressure formula; then the air density and mass flow rate are calculated in combination with the gas constant; finally, a complete internal flow model is constructed through the flow parameter distribution to describe the basic characteristics of the inlet duct outlet airflow.

[0139] S2-2, construct the low-speed internal flow model of the inlet;

[0140] In step S2-2, a low-speed internal flow model of the air inlet is constructed, including:

[0141] S2-2-1. When the flight Mach number Ma is less than 0.6, the total pressure recovery coefficient is calculated using the empirical formula;

[0142] σ=0.782+0.147cos(Ma)+0.152sin(Ma);

[0143] Where σ is the total pressure recovery coefficient.

[0144] When the flight Mach number Ma<0.6, the inlet exhibits flow characteristics similar to a large trumpet-shaped suction intake device. At this time, the inlet outlet flow coefficient q(λ2) is usually greater than 1, and the inlet can effectively capture and compress the air flow. The total pressure recovery coefficient σ calculated by the empirical formula can accurately describe the aerodynamic performance of the inlet under low-speed conditions, ensuring the adaptability and accuracy of the model at different flight speeds.

[0145] This empirical formula is based on a large amount of experimental data and can accurately predict the total pressure recovery performance of the inlet under low-speed flight conditions. Through this step, the low-speed internal flow model of the inlet can reflect the dynamic changes of the airflow under low Mach numbers, providing the necessary data support for the accurate coupling of the comprehensive model.

[0146] S2-3. Construct an inlet outflow model.

[0147] Specifically, the inlet modeling is divided into internal flow, low-speed internal flow and external flow models. Each part takes into account important characteristics such as the inlet outlet parameters, characteristics under low-speed conditions and external flow resistance, ensuring that the model can cover the aerodynamic characteristics of the entire flight envelope.

[0148] In step S2-3, an inlet outflow model is constructed, including:

[0149] S2-3-1. Obtain the overflow resistance, venting resistance and boundary layer discharge resistance of the air inlet;

[0150] S2-3-2, inputting the overflow resistance, the deflation resistance and the boundary layer discharge resistance of the inlet duct, and calculating the total outflow resistance;

[0151] The expression for calculating the sum of the outflow resistance is: in =D sp +D bp +D bl

[0152] Among them, D sp is the overflow resistance, which describes the overflow characteristics of the excess air in the intake duct; D bp The deflation resistance describes the resistance of the external airflow caused by the deflation characteristics; D blThe boundary layer shedding resistance describes the resistance caused by the shedding effect of the boundary layer.

[0153] The inlet involved in this embodiment is a supersonic dual inlet, and its outflow characteristics are the changes of the components of various forces on its outer surface along the flight direction with the flight Mach number and flow coefficient. Its outflow resistance is composed of overflow resistance, venting resistance and boundary layer discharge resistance.

[0154] S2-3-3, obtain the total external flow resistance, flight Mach number, inlet cross-sectional area, and specific heat ratio of air;

[0155] S2-3-4. Calculate the inlet drag coefficient by inputting the total outflow resistance, flight Mach number, inlet cross-sectional area, and specific heat ratio of air;

[0156] The expression for calculating the inlet resistance coefficient is:

[0157] Among them, C in is the inlet resistance coefficient, D in is the total outflow resistance, γ is the specific heat ratio of air, A c is the inlet cross-sectional area, p0 is the total pressure of the air;

[0158] S2-3-5. Obtain the drag coefficient, flight speed, air density and cross-sectional area;

[0159] S2-3-6. Calculate the total external resistance of the air inlet and tail nozzle according to the drag coefficient, flight speed, air density and cross-sectional area;

[0160] The expression for calculating the total external resistance of the air inlet and the tail nozzle is:

[0161] Among them, F inlet,drag represents the total external resistance of the air inlet and tail nozzle, ρ is the air density, v is the flight speed, A c is the cross-sectional area of ​​the inlet duct.

[0162] Specifically, the external flow model calculates the total external flow resistance and drag coefficient by calculating the overflow resistance, outflow resistance and boundary layer resistance. Finally, the external resistance is calculated by combining the flight speed and air density to provide input for the thrust balance calculation of the entire inlet model.

[0163] The modeling method further comprises step S3:

[0164] S3. In the initial model, construct an engine component level model;

[0165] In step S3, in the initial model, an engine component level model is constructed, including:

[0166] S3-1, obtaining the converted flow rate at the inlet duct outlet and the converted flow rate at the engine fan, and constructing a residual equation for matching and balancing the inlet duct and engine flow rates;

[0167] The residual equation for the balance between the intake duct and the engine flow is:

[0168] Among them, m 2,c Indicates the calculated flow rate at the inlet duct outlet, m 21,c represents the calculated flow rate of the engine fan, ∈1 represents the residual between the intake duct and the engine flow rate;

[0169] S3-2, obtaining the high-pressure turbine guide valve inlet flow rate and the high-pressure turbine characteristic line to calculate the flow rate, and constructing the residual equation of the high-pressure turbine rotor inlet flow balance;

[0170] The residual equation for the high-pressure turbine rotor inlet flow balance is:

[0171] Among them, m 41,cx represents the calculated inlet flow rate of the high-pressure turbine guide, m 41,c represents the actual inlet flow of the high-pressure turbine, ∈2 represents the residual between the inlet flow of the high-pressure turbine guide valve and the flow calculated by the high-pressure turbine characteristic line;

[0172] S3-3, obtaining the flow rate of the low-pressure turbine guide vane inlet and the flow rate calculated by the high-pressure turbine characteristic line, and constructing the residual equation of the low-pressure turbine rotor inlet flow balance;

[0173] The residual equation for the low-pressure turbine rotor inlet flow balance is:

[0174] Among them, m 45,cx represents the calculated inlet flow rate of the low-pressure turbine guide vane, m 45,c represents the actual inlet flow of the low-pressure turbine, ∈3 represents the residual between the inlet flow of the low-pressure turbine guide valve and the flow calculated by the high-pressure turbine characteristic line;

[0175] S3-4, obtaining the static pressure at the inner and outer outlets of the engine, and constructing the residual equation of the static pressure balance at the inlet of the mixing chamber;

[0176] The residual equation of static pressure balance at the mixing chamber inlet is:

[0177] Among them, p 16,s represents the static pressure at the outlet of the vortex mixing chamber in the engine, p 6,s represents the static pressure of the outer vortex mixing chamber, ∈4 represents the residual between the static pressure at the engine outlet and the static pressure at the outer vortex outlet;

[0178] S3-5, obtain the flow rate of the tail nozzle and the flow characteristics of the tail nozzle to calculate the flow rate, and construct the residual equation of the tail nozzle throat flow balance;

[0179] The residual equation of the tail nozzle throat flow balance is:

[0180] Among them, m 8,x represents the calculated flow rate of the tail nozzle, m8 represents the actual flow rate of the tail nozzle, ∈5 represents the residual between the flow rate of the tail nozzle and the flow rate calculated by the internal flow characteristics of the tail nozzle;

[0181] S3-6, obtaining the power consumed by the fan and the power provided by the low-pressure turbine, and constructing a residual equation for the power balance of the low-pressure rotor;

[0182] The residual equation for the low-pressure rotor power balance is:

[0183] Among them, W f Indicates the power consumed by the fan, W lt represents the power provided by the low-pressure turbine, η represents the mechanical transmission efficiency, ∈6 represents the residual between the power consumed by the fan and the power provided by the low-pressure turbine;

[0184] S3-7, obtaining the power consumed by the fan and the power provided by the high-pressure turbine, and constructing a residual equation balanced by the power of the high-pressure rotor;

[0185] The residual equation for the high-pressure rotor power balance is:

[0186] Among them, W c represents the power consumed by the compressor, Wht represents the power provided by the high-pressure turbine, η h represents the mechanical transmission efficiency of the high-pressure rotor shaft, ∈7 represents the residual between the power consumed by the fan and the power provided by the high-pressure turbine

[0187] S3-8, combining the residual variances from S3-1 to S3-7 to obtain the collaborative engineering equations;

[0188] S3-9. Use the dynamic characteristic equation to calculate the speed change rate of the high and low pressure rotors under dynamic conditions;

[0189]

[0190] Among them, N f and N c are the speeds of the fan and compressor respectively, N lt and N ht are the speeds of the low-pressure turbine and the high-pressure turbine, respectively, and J f and J c are the rotational inertia of the fan and compressor respectively;

[0191] S3-10, combining the collaborative engineering group with the dynamic characteristic equation to form the engine component level model

[0192] Specifically, the engine component-level model starts with key factors such as flow matching, static pressure balance and power balance, and gradually builds a complete collaborative model including components such as fans, turbines, and tail nozzles, and describes their working characteristics through residual equations and dynamic characteristic equations.

[0193] The modeling method further comprises step S4:

[0194] S4, treating the combination of the intake duct component-level model and the engine component-level model as a comprehensive dynamic model, and solving the comprehensive dynamic model using a collaborative working equation group;

[0195] In step S4, the integrated dynamic model is solved using the collaborative working equations, including:

[0196] S4-1, combining the intake duct component level model and the engine component level model to construct a comprehensive dynamic model;

[0197] S4-2, selecting initialization parameters to perform iterative updates on the comprehensive dynamic model;

[0198] S4-3, when the norm of the residual matrix is ​​updated iteratively to meet the convergence threshold, the output balance of the comprehensive dynamic model is determined, otherwise the iteration continues;

[0199] ∈ n =[∈1,∈2,…,∈7] T

[0200] ||∈ n ||≤δ

[0201] Among them, ∈ n represents the residual matrix, ||∈ n || represents the norm of the residual matrix, and δ represents the convergence threshold.

[0202] Specifically, a complete dynamic integrated model is constructed by combining the intake duct and the engine model, and the residual matrix is ​​updated during the iterative solution process to ensure that the output of the model reaches a balanced state, providing accurate initial values ​​for subsequent optimization.

[0203] Embodiment 2: The technical solution of Embodiment 2 is different from that of Embodiment 1 in that, in step S5, the trust region Newton one-pass algorithm is used to optimize the cooperative working equations to update the guessed parameters of the comprehensive dynamic model; wherein the guessed parameters include: inlet flow coefficient, fan pressure ratio coefficient, compressor pressure ratio coefficient, high-pressure turbine inlet flow, and low-pressure turbine inlet flow. Including:

[0204] Using the Trust Region Newton One-Pass Algorithm, the objective function of the collaborative engineering equations is defined;

[0205] The trust region Newton one-pass algorithm passes through the iteration point x k An approximate objective optimization function of the objective function f(x) is constructed to solve the problem;

[0206] Among them, the approximate objective optimization function is:

[0207]

[0208] Among them, s k =x k+1 -x k ;

[0209] s k is the step size, which means from the current iteration point x k To the next iteration point x k+1 The displacement is used to optimize the update of variables; x k is the current iteration point, indicating the location of the current optimization solution, x k+1 is the next iteration point, indicating the new solution updated after this iteration; the step definition formula describes the step length between the current solution and the next solution during the iteration process, and the step length s k It is the core of variable update.

[0210] F k (s k ) is the approximate objective function, which is constructed by quadratic approximation of f(x) and is used to find the optimal solution in the current trust region;

[0211] f(x k ) is the objective function f(x) at the current iteration point x k The function value of ;

[0212] g k is the objective function f(x) at the iteration point x k The gradient vector at is defined as:

[0213]

[0214] The gradient vector describes the direction and rate of change of the objective function at the current iteration point; is the transpose of the gradient vector, which changes the gradient vector from a column vector to a row vector for use with the step size s k Perform inner product operation; is a linear term, indicating that the objective function is at the current point x k Along step length s k The first-order change in direction; H k is the objective function f(x) at x kThe Hessian matrix at is defined as the second-order partial derivative matrix:

[0215]

[0216] The Hessian matrix reflects the second-order curvature characteristics of the objective function; is a quadratic term used to adjust the weight of the quadratic approximation model;

[0217] Among them, the approximate objective function formula describes the change of the objective function through the gradient and Hessian matrix and the quadratic approximation term, which is the core model of the trust region Newton one-pass algorithm;

[0218] Step constraint formula: st||s k ||≤h k ||s k || is the step length s k The second norm of is defined as: It represents the length of the step; h k is the trust region radius, which indicates the upper bound of the step size in the current iteration; it is used to limit the step size s k The size of , thus ensuring that the optimization solution converges within the trust region; ||s k ||≤h k As a constraint condition, the step length s is required k The length of the trust region does not exceed the radius h k ; By limiting the step size, the optimization process is ensured to be carried out within the trust region, thus ensuring convergence and stability.

[0219] Specifically, the Trust Region Newton One-Pass Algorithm constructs the objective function, uses the gradient and Hessian matrix for quadratic approximation, and combines step size restrictions to optimize the key parameters in the dynamic model. Finally, through iterative updates, the output of the intake and engine models is adjusted to the optimal state.

[0220] As an implementation method of this embodiment, the trust region is defined as a given trust region radius h k Inlet flow coefficient φ in,k Neighborhood of;

[0221] The definition of the trust region is:

[0222]

[0223] Among them, Ω k represents the trust region in the kth iteration; the trust region is defined as the optimization variable x restricted to the iteration point φ in,k The area within a certain range nearby is determined by the trust region radius h k Decide;

[0224] x represents the possible variable value (a vector) in the optimization process, i.e., the candidate solution;

[0225] φ in,k represents the inlet flow coefficient at the current iteration point (a vector), which is the solution of the trust region Newton one-pass algorithm at the kth iteration;

[0226] ||x-φ in,k || represents vector x and φ in,k The distance between is defined using the binorm (Euclidean distance): Among them, x i and φ in,k are vectors x and φ respectively in,k The i-th component of ;

[0227] Trust region radius h k is a scalar value that limits the variable x to the current iteration φ in,k The maximum distance ensures that the optimization process is only performed within the trust region.

[0228] Specifically, the trust region ensures that the optimization process is carried out in a local range by limiting the step size and radius, thereby improving the convergence of the algorithm and avoiding model instability caused by large step size updates. The neighborhood is defined as the Euclidean distance of the variable, which describes the range of parameter adjustment during the iteration process.

[0229] Specifically, the core of this method lies in comprehensive modeling and collaborative optimization. The parameter initialization of S1 is the basis for subsequent models; S2 and S3 respectively establish component-level models of the intake duct and the engine; S4 combines these two parts into a dynamic comprehensive model and uses collaborative equations to solve its steady-state or dynamic characteristics; finally, the optimization algorithm is introduced in S5 to further adjust the model parameters and improve the overall performance of the system.

[0230] The above embodiments can be implemented in whole or in part by software, hardware, firmware or other arbitrary combinations. When implemented using software, the above embodiments can be implemented in whole or in part in the form of a computer program product. The computer program product includes one or more computer instructions or computer programs. When the computer instructions or computer programs are loaded or executed on a computer, the process or function described in the embodiment of the present application is generated in whole or in part. The computer can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable devices. The computer instructions can be stored in a computer-readable storage medium, or transmitted from one computer-readable storage medium to another computer-readable storage medium, for example, the computer instructions can be transmitted from a website site, computer, server or data center by wired (e.g., infrared, wireless, microwave, etc.) mode to another website site, computer, server or data center.

[0231] The computer-readable storage medium may be any available medium that can be accessed by a computer or a data storage device such as a server or a data center that includes one or more available media. The available medium may be a magnetic medium (e.g., a floppy disk, a hard disk, a tape), an optical medium (e.g., a DVD), or a semiconductor medium. The semiconductor medium may be a solid-state hard disk.

[0232] In the several embodiments provided in the present application, it should be understood that the disclosed systems, devices and methods can be implemented in other ways. For example, the device embodiments described above are only schematic. For example, the division of the units is only a logical function division of a waterway underwater terrain change analysis system and method. There may be other division methods in actual implementation. For example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be through some interfaces, indirect coupling or communication connection of devices or units, which can be electrical, mechanical or other forms.

[0233] The above description is only a specific implementation manner of the present application, but the protection scope of the present application is not limited thereto. Any technician familiar with the technical field can easily think of changes or substitutions within the technical scope disclosed in the present application, which should be included in the protection scope of the present application.

Claims

1. A component-level supersonic inlet comprehensive model modeling method based on an aircraft engine, characterized in that: include: S1, obtain initialization parameters to define the initial model; S2. In the initial model, construct an intake duct component level model; S3. In the initial model, construct an engine component level model; S4, treating the combination of the intake duct component-level model and the engine component-level model as a comprehensive dynamic model, and solving the comprehensive dynamic model using a collaborative working equation group; S5. Optimizing the collaborative working equations using the trust region Newton one-pass algorithm to update the guessed parameters of the comprehensive dynamic model; The guessed parameters include: inlet flow coefficient, fan pressure ratio coefficient, compressor pressure ratio coefficient, high-pressure turbine inlet flow and low-pressure turbine inlet flow.

2. The component-level supersonic inlet comprehensive modeling method based on an aircraft engine according to claim 1 is characterized in that: The steps of defining the initial model include: S1-1, obtaining initialization parameters, including: flight altitude H, flight Mach number Ma, fuel quantity, nozzle outlet area, fan pressure ratio coefficient, compressor pressure ratio coefficient; S2. Define the initial state of the component-level model according to the initialization parameters.

3. The component-level supersonic inlet comprehensive modeling method based on an aircraft engine according to claim 2 is characterized in that: The steps for building the intake duct component-level model include: S2-1, construct the inlet flow model; S2-2, construct the low-speed internal flow model of the inlet; S2-3. Construct an inlet outflow model.

4. The component-level supersonic inlet comprehensive modeling method based on an aircraft engine according to claim 2 is characterized in that: The steps to construct the inlet flow model include: S2-1-1. Obtain the total temperature and flight Mach number in the flight state, and substitute them into the following formula to generate the total temperature at the inlet outlet; Among them, T0 is the total temperature in flight, Ma is the flight Mach number, and κ is the air adiabatic index; S2-1-2. Obtain the total pressure and flight Mach number in the flight state, substitute them into the following formula to generate the total pressure at the inlet outlet; Among them, p0 is the total pressure in flight state, and p2 is the total pressure at the inlet outlet; S2-1-3, obtaining the total temperature and total pressure at the inlet duct outlet, and generating the air density at the inlet duct outlet; The expression for calculating the air density at the inlet duct outlet is: Among them, ρ2 is the air density at the inlet outlet, and R is the gas constant; S2-1-4. Obtain the outlet cross-sectional area, velocity coefficient, and flow function, and substitute them into the following formula to calculate the air mass flow rate; The expression for calculating air mass flow is: in, is the air mass flow rate at the inlet duct outlet, A2 is the cross-sectional area of ​​the inlet duct outlet, λ2 is the inlet duct outlet velocity coefficient, and q(λ2) is the inlet duct outlet flow coefficient; S2-1-5. Based on the calculated total temperature at the inlet duct outlet, total pressure at the inlet duct outlet and air mass flow rate, establish a flow parameter distribution to construct the inlet duct internal flow model.

5. The component-level supersonic inlet comprehensive modeling method based on an aircraft engine according to claim 1 is characterized in that: Construct a low-speed internal flow model for the inlet, including: S2-2-1. When the flight Mach number Ma is less than 0.6, the total pressure recovery coefficient is calculated using the empirical formula; σ=0.782+0.147cos(Ma)+0.152sin(Ma); Where σ is the total pressure recovery coefficient.

6. The component-level supersonic inlet comprehensive modeling method based on an aircraft engine according to claim 1 is characterized in that: The steps of building the inlet outflow model include: S2-3-1. Obtain the overflow resistance, venting resistance and boundary layer discharge resistance of the air inlet; S2-3-2, inputting the overflow resistance, the deflation resistance and the boundary layer discharge resistance of the inlet duct, and calculating the total outflow resistance; The expression for calculating the sum of the outflow resistance is: in =D sp +D bp +D bl Among them, D sp is the overflow resistance, which describes the overflow characteristics of the excess air in the intake duct; D bp The deflation resistance describes the resistance of the external airflow caused by the deflation characteristics; D bl The boundary layer shedding resistance describes the resistance caused by the shedding effect of the boundary layer. S2-3-3, obtain the total external flow resistance, flight Mach number, inlet cross-sectional area, and specific heat ratio of air; S2-3-4. Calculate the inlet drag coefficient by inputting the total outflow resistance, flight Mach number, inlet cross-sectional area, and specific heat ratio of air; The expression for calculating the inlet resistance coefficient is: Among them, C in is the inlet resistance coefficient, D in is the total outflow resistance, γ is the specific heat ratio of air, A c is the inlet cross-sectional area, p0 is the total pressure of the air; S2-3-5. Obtain the drag coefficient, flight speed, air density and cross-sectional area; S2-3-6. Calculate the total external resistance of the air inlet and tail nozzle according to the drag coefficient, flight speed, air density and cross-sectional area; The expression for calculating the total external resistance of the air inlet and the tail nozzle is: Among them, F inlet,drag represents the total external resistance of the air inlet and tail nozzle, ρ is the air density, v is the flight speed, A c is the cross-sectional area of ​​the inlet duct.

7. The method for modeling a component-level supersonic inlet comprehensive model based on an aircraft engine according to claim 1, characterized in that: The steps for building the engine component-level model include: S3-1, obtaining the converted flow rate at the inlet duct outlet and the converted flow rate at the engine fan, and constructing a residual equation for matching and balancing the inlet duct and engine flow rates; The residual equation for the balance between the intake duct and the engine flow is: Among them, m 2,c Indicates the calculated flow rate at the inlet duct outlet, m 21,c represents the calculated flow rate of the engine fan, ∈1 represents the residual between the intake duct and the engine flow rate; S3-2, obtaining the high-pressure turbine guide valve inlet flow rate and the high-pressure turbine characteristic line to calculate the flow rate, and constructing the residual equation of the high-pressure turbine rotor inlet flow balance; The residual equation for the high-pressure turbine rotor inlet flow balance is: Among them, m 41,cx represents the calculated inlet flow rate of the high-pressure turbine guide, m 41,c represents the actual inlet flow of the high-pressure turbine, ∈2 represents the residual between the inlet flow of the high-pressure turbine guide valve and the flow calculated by the high-pressure turbine characteristic line; S3-3, obtaining the flow rate of the low-pressure turbine guide vane inlet and the flow rate calculated by the high-pressure turbine characteristic line, and constructing the residual equation of the low-pressure turbine rotor inlet flow balance; The residual equation for the low-pressure turbine rotor inlet flow balance is: Among them, m 45,cx represents the calculated inlet flow rate of the low-pressure turbine guide vane, m 45,c represents the actual inlet flow of the low-pressure turbine, ∈3 represents the residual between the inlet flow of the low-pressure turbine guide valve and the flow calculated by the high-pressure turbine characteristic line; S3-4, obtaining the static pressure at the inner and outer outlets of the engine, and constructing the residual equation of the static pressure balance at the inlet of the mixing chamber; The residual equation of static pressure balance at the mixing chamber inlet is: Among them, p 16,s represents the static pressure at the outlet of the vortex mixing chamber in the engine, p 6，s represents the static pressure of the outer vortex mixing chamber, ∈4 represents the residual between the static pressure at the engine outlet and the static pressure at the outer vortex outlet; S3-5, obtain the flow rate of the tail nozzle and the flow characteristics of the tail nozzle to calculate the flow rate, and construct the residual equation of the tail nozzle throat flow balance; The residual equation of the tail nozzle throat flow balance is: Among them, m 8,x represents the calculated flow rate of the tail nozzle, m8 represents the actual flow rate of the tail nozzle, ∈5 represents the residual between the flow rate of the tail nozzle and the flow rate calculated by the internal flow characteristics of the tail nozzle; S3-6, obtaining the power consumed by the fan and the power provided by the low-pressure turbine, and constructing a residual equation for the power balance of the low-pressure rotor; The residual equation for the low-pressure rotor power balance is: Among them, W f Indicates the power consumed by the fan, W lt represents the power provided by the low-pressure turbine, η represents the mechanical transmission efficiency, ∈6 represents the residual between the power consumed by the fan and the power provided by the low-pressure turbine; S3-7, obtaining the power consumed by the fan and the power provided by the high-pressure turbine, and constructing a residual equation balanced by the power of the high-pressure rotor; The residual equation for the high-pressure rotor power balance is: Among them, W c represents the power consumed by the compressor, Wht represents the power provided by the high-pressure turbine, η h represents the mechanical transmission efficiency of the high-pressure rotor shaft, ∈7 represents the residual between the power consumed by the fan and the power provided by the high-pressure turbine S3-8, combining the residual variances from S3-1 to S3-7 to obtain the collaborative engineering equations; S3-9. Use the dynamic characteristic equation to calculate the speed change rate of the high and low pressure rotors under dynamic conditions; Among them, N f and N c are the speeds of the fan and compressor respectively, N lt and N ht are the speeds of the low-pressure turbine and the high-pressure turbine, respectively, and J f and J c are the rotational inertia of the fan and compressor respectively; S3-10. Combining the collaborative working engineering group with the dynamic characteristic equation to form the engine component level model.

8. The method for modeling a component-level supersonic inlet comprehensive model based on an aircraft engine according to claim 1, characterized in that: The method of using the collaborative working equation to solve the intake duct component level model and the engine component level model includes: S4-1, combining the intake duct component level model and the engine component level model to construct a comprehensive dynamic model; S4-2, selecting initialization parameters to perform iterative updates on the comprehensive dynamic model; S4-3, when the norm of the residual matrix is ​​updated iteratively to meet the convergence threshold, the output balance of the comprehensive dynamic model is determined, otherwise the iteration continues; ∈ n ＝[∈1，∈2，…，∈7] T ||∈ n ||≤δ Among them, ∈ n represents the residual matrix, ||∈ n || represents the norm of the residual matrix, and δ represents the convergence threshold.

9. The method for modeling a component-level supersonic inlet comprehensive model based on an aircraft engine according to claim 1, characterized in that: The collaborative set of equations is optimized using a trust-region Newton one-pass algorithm to update the guessed parameters of the comprehensive dynamic model, including: Using the Trust Region Newton One-Pass Algorithm, the objective function of the collaborative engineering equations is defined; The trust region Newton one-pass algorithm passes through the iteration point x k An approximate objective optimization function of the objective function f(x) is constructed to solve the problem; Among them, the approximate objective optimization function is: Among them, s k =X k+1 -X k ; s k is the step length, which means from the current iteration point x k To the next iteration point x k+1 The displacement is used to optimize the update of variables; x k is the current iteration point, indicating the location of the current optimization solution, X k+1 is the next iteration point, indicating the new solution updated after this iteration; the step definition formula describes the step length between the current solution and the next solution during the iteration process, and the step length s k It is the core of variable update; F k (s k ) is the approximate objective function, which is constructed by quadratic approximation f(x) and is used to find the optimal solution in the current trust region; f(x k ) is the objective function f(x) at the current iteration point x k The function value of g k is the objective function f(x) at the iteration point x k The gradient vector at is defined as: The gradient vector describes the direction and rate of change of the objective function at the current iteration point; is the transpose of the gradient vector, which changes the gradient vector from a column vector to a row vector for use with the step size s k Perform inner product operation; is a linear term, indicating that the objective function is at the current point x k Along step length s k The first-order change in direction; H k is the objective function f(x) at x k The Hessian matrix at is defined as the second-order partial derivative matrix: is a quadratic term used to adjust the weight of the quadratic approximation model; ||s k || is the step length s k The second norm of is defined as: It represents the length of the step; h k is the trust region radius, which indicates the upper bound of the step size in the current iteration; it is used to limit the step size s k The size of , thus ensuring that the optimization solution converges within the trust region; ||s k ||≤h k As a constraint condition, the step length s is required k The length of the trust region does not exceed the radius h k .

10. The method for modeling a component-level supersonic inlet comprehensive model based on an aircraft engine according to claim 9, characterized in that: The trust region is defined as a given trust region radius h k Inlet flow coefficient φ in，k The definition of the trust region is: Among them, Ω k represents the trust region in the kth iteration; the trust region is defined as the optimization variable x restricted to the iteration point φ in，k The area within a certain range nearby is determined by the trust region radius h k Decide; x represents the possible variable value in the optimization process; φ in，k represents the inlet flow coefficient at the current iteration point, which is the solution of the trust region Newton one-pass algorithm in the kth iteration; ||x-φ in，k || represents vector x and φ in，k The distance between is defined using the two-norm: Among them, x i and φ in，k are vectors x and φ respectively in，k The i-th component of .

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