Large single-chamber double-thrust solid engine performance optimization method based on multi-objective optimization

Abstract

The invention provides a large-scale single-chamber double-thrust solid engine performance optimization method based on multi-objective optimization, which comprises the following steps: designing an initial scheme of an engine, constructing an optimization model, determining optimization objectives and optimizing the optimization objectives respectively, evaluating and determining a final optimization scheme, obtaining a candidate solution set according to single-objective optimization, and calculating the performance of the large-scale single-chamber double-thrust solid engine according to the candidate solution set. The optimal value is determined by adopting analysis and evaluation methods such as analytic hierarchy process and weighted analysis, the multi-target optimization design of the performance of the large-size single-chamber double-thrust solid engine is achieved through the scheme design of the high-performance single-chamber double-thrust engine, multi-target optimization is carried out according to an optimization model and engine design constraint conditions, and the performance of the large-size single-chamber double-thrust solid engine is optimized. And the design of the large-size single-chamber double-thrust solid engine under the constraint condition is determined. The mass ratio of the engine and the range of the aircraft are improved under the condition that the appearance size of the engine is the same and the weight is basically unchanged.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft engine design, in particular to a method for optimizing the performance of a large single-chamber dual-thrust solid engine. Background Art

[0002] Solid rocket engines are widely used in missiles and rockets. Large-sized single-chamber dual-thrust solid engines are the preferred power solution for medium- and long-range rocket engines, which helps to improve the overall performance of rockets. The engine design solution is related to the improvement of rocket launch performance, flight stability and range, and is a difficult and key issue in the design of medium and large-caliber engines.

[0003] At present, the research on engine optimization design methods at home and abroad is mostly carried out on single variable optimization such as engine mass ratio and total impulse. There are few studies on multi-objective comprehensive optimization methods such as the working characteristics of the active section of the rocket. There is no multi-objective optimization method for the performance of large-scale single-chamber dual-thrust solid engine. Summary of the invention

[0004] In order to overcome the shortcomings of the prior art, the present invention provides a large-scale single-chamber dual-thrust solid engine performance optimization method based on multi-objective optimization, solves the multi-objective optimization design of large-size single-chamber dual-thrust solid engine performance, and carries out multi-objective optimization according to the optimization model and engine design constraints to determine the design of a large-size single-chamber dual-thrust solid engine under the constraints. The present invention includes four steps: initial engine scheme design, construction of an optimization model, determination of optimization targets and optimization of each, evaluation and determination of the final optimization scheme, obtaining a candidate solution set according to single-objective optimization, and using analytical evaluation methods such as hierarchical analysis and weighted analysis to determine the optimal value, so as to achieve high-performance single-chamber dual-thrust engine scheme design.

[0005] The technical solution adopted by the present invention to solve the technical problem includes the following steps:

[0006] Step 1: Initial engine design: Design the engine's overall dimensions, structure, charge and internal ballistics;

[0007] Complete the design of charge type, propellant performance and internal ballistic scheme according to the total engine impulse and working time requirements;

[0008] Complete the combustion chamber casing design according to the engine operating pressure, size and interface requirements;

[0009] According to the internal ballistic design results and working time requirements, the nozzle assembly and thermal protection scheme design are completed. The initial shape scheme of the engine is as follows: Figure 2 As shown;

[0010] Step 2: Construct an optimization model: This is used to establish an optimization algorithm under the constraints of performance, size, and weight. According to the optimization goals of the highest mass ratio, the highest total impulse, and the longest external trajectory matching range, a linear weighted algorithm is used in combination with the optimization weights.

[0011] Step 3: Determine the optimization objectives and optimize them separately: obtain the single variable optimization candidate solution set of multiple objectives, the charge scheme and propellant performance selection set, throat diameter selection and expansion ratio scheme set that meet the maximum total impulse condition;

[0012] Step 4: Evaluate and determine the final optimization plan: Analyze the weights of each optimization goal, and conduct comprehensive analysis and evaluation to obtain the final optimization design plan that is both objective and reasonable.

[0013] The specific steps of step 1 are:

[0014] Step 1.1: Charge and interior ballistics design;

[0015] The charge design should improve the charge density while meeting the overall ballistic index requirements for the engine. On the one hand, mature charge type solutions such as tube slots, wing columns, and star holes should be selected to reduce the process complexity and development cost of the charge. On the other hand, the filling coefficient should be selected to improve the charge's adaptability to a wide temperature environment. On this basis, the charge column structure and propellant performance should be optimized to determine the charge type solution.

[0016] After the charge design is completed, the internal ballistics is estimated according to the principle of parallel layer displacement and calculated according to the zero-dimensional steady-state engine internal ballistic differential equation. The calculation formula is as follows:

[0017]

[0018] Where: - gas mass generation rate (kg / s); Ka - burning rate correction factor; a - burning rate at calculated temperature; P C - Combustion chamber pressure, n-pressure index; A b - charge burning area; ρ p - propellant density; σ P - Burning rate temperature sensitivity coefficient (1 / K); T - calculation temperature; T0 - reference temperature; η c - Characteristic velocity coefficient of the propellant; R g - Gas constant of fuel gas; T g - Combustion temperature of the gas; Vc - Free volume of the combustion chamber cavity; η P - is the total pressure recovery coefficient of the nozzle flow; A t - throat area; C* - characteristic velocity of propellant m / s; - Nozzle flow correction factor;

[0019] According to formula (1) and formula (2), combined with the performance parameters of the grain and propellant, the combustion chamber working pressure P is calculated as C , throat area A t , expansion ratio ξ, working time, two-stage working time ratio, total engine impulse and two-stage impulse ratio are used as basic parameters for optimization design;

[0020] Step 1.2: Shell design;

[0021] Use the fourth strength theory to calculate the shell wall thickness and determine the combustion chamber shell wall thickness δ min :

[0022]

[0023] in, is the pressure fluctuation coefficient; ξ is the weld strength coefficient; D i is the inner diameter of the cylinder; P max is the maximum working pressure; [σ] is the allowable stress of the material, [σ] = σb / nb, the engine shell wall thickness δ min ,At the same time, according to the shape and weight requirements, the length and detailed structure design of the combustion chamber shell are completed, and the shell weight design parameters are obtained as the basic parameters for optimization design;

[0024] Step 1.3: Nozzle Design

[0025] Based on the calculated throat area A t and expansion ratio, nozzle design completes the nozzle's aerodynamic design and thermal protection design;

[0026] The aerodynamic design is the shape parameters of the convergent section, throat and diffuser section, which makes the nozzle more efficient. The aerodynamic design of the nozzle ensures the continuity of the profile, and adopts the form of tangent connection between the inlet arc and the straight line to avoid the existence of compression waves and shock waves caused by the discontinuity of the profile. The downstream design of the throat is the expansion half angle selection of the expansion section to reduce the sudden expansion loss caused by the profile, and complete the internal profile design of the nozzle under the conditions of meeting the throat diameter, outlet diameter and external size constraints.

[0027] The throat diameter is:

[0028]

[0029] Outlet diameter:

[0030] D e =D t ξ 0.5

[0031] Combined with the designed engine operating pressure and working time, the design of the nozzle thermal protection structure is completed, and the design parameters of the nozzle weight, length and expansion half angle are obtained as the basic parameters for optimization design;

[0032] The specific steps of constructing the optimization model in step 2 are:

[0033] Under the constraints of performance and size, three optimization objectives, namely engine mass ratio, total impulse and aircraft range, are established in combination with design and use requirements, and the least squares loss function of each optimization objective is determined so that each independent optimization objective converges to the optimal parameter.

[0034] The research object is a large-size single-chamber dual-thrust solid engine scheme, or a single-chamber large-thrust engine scheme. The optimization parameters include the engine combustion chamber pressure P obtained by the previous calculation. C , throat area A t 、Combustion chamber shell wall thickness δ min , expansion ratio ξ, working time t, two-stage working time ratio, total engine impulse and two-stage impulse ratio, nozzle weight, nozzle length and expansion half angle;

[0035] The linear weighted method is used to optimize the aircraft's longest range and weight it, which can be expressed as:

[0036]

[0037] stg(x)≥0,i∈[1,M]

[0038] h j (x)=0,j∈[1,L]

[0039] Through the linear weighted algorithm, multi-objective optimization is transformed into a more achievable single-objective optimization problem model.

[0040] The specific steps of step 3 are:

[0041] The optimization target weights are set differently. When the range optimization is the main target, the range optimization weight is set to 0.6, the mass ratio and total impulse are set to 0.2, and the main optimization target is optimized. Under the constraints of engine working pressure and working time, the best result is selected as the candidate solution set to obtain multiple candidate solutions.

[0042] In step 4, the steps of evaluating and determining the final optimization solution are:

[0043] Evaluate the main objective function value of the candidate target solutions obtained, and comprehensively analyze and evaluate the engineering feasibility of the engine charge and internal ballistics solution, shell solution, and nozzle solution under the optimal conditions of the main target (range), so as to obtain the final solution that meets the optimization expectations, meets the usage requirements, and has the ability to be engineered.

[0044] The beneficial effect of the present invention is that the performance optimization method of a large-scale single-chamber dual-thrust solid engine based on multi-objective optimization provided by the present invention solves the multi-objective optimization design problem of solid power for medium- and long-range rockets. The optimization design of a large-scale single-chamber dual-thrust solid engine is achieved under multiple constraints, significantly improving the overall performance of the engine and rocket. A large-scale engine adopts a metal structure, composite propellant charge, and a fixed composite nozzle scheme. During the demonstration process, this optimization method was used to achieve an improvement in the engine mass ratio and the range of the aircraft while ensuring that the engine's overall dimensions are the same and the weight is basically unchanged. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Figure 1 It is a flow chart of a method for optimizing the performance of a large-sized single-chamber dual-thrust solid engine based on multi-objective optimization in an embodiment of the present invention.

[0046] Figure 2 It is a schematic diagram of the initial design scheme of the engine according to the embodiment of the present invention.

[0047] Figure 3 It is a schematic diagram comparing the improvement of the overall performance of the engine and the aircraft after optimization using the present invention. DETAILED DESCRIPTION

[0048] The present invention is further described below in conjunction with the accompanying drawings and embodiments.

[0049] The flow chart of the method for optimizing the performance of a large single-chamber dual-thrust solid rocket engine based on multi-objective optimization in this embodiment is as follows: Figure 1 As shown, it mainly includes four steps: initial engine scheme design, building an optimization model, determining the optimization goals and optimizing each, and evaluating and determining the final optimization scheme.

[0050] Step 1: Charge and interior ballistics design;

[0051] The charge design should maximize the charge density while meeting the overall ballistic index requirements for the engine. On the one hand, mature charge types such as tube slots, wing columns, and star holes should be selected as much as possible to reduce the process complexity and development cost of the charge. On the other hand, an appropriate loading coefficient should be selected to improve the charge's adaptability to a wide temperature environment. On this basis, the charge column structure and propellant performance should be optimized to determine the charge type.

[0052] After the charge design is completed, the internal ballistics is estimated according to the principle of parallel layer displacement and calculated according to the zero-dimensional steady-state engine internal ballistic differential equation. The calculation formula is as follows:

[0053]

[0054] Where: - gas mass generation rate (kg / s); Ka - burning rate correction factor; a - burning rate at calculated temperature; P C - combustion chamber pressure, n-pressure index; Ab-charge combustion area; ρp-propellant density; σ P - Burning rate temperature sensitivity coefficient (1 / K); T - calculation temperature; T0 - reference temperature; η c - characteristic velocity coefficient of propellant; Rg - gas constant of fuel gas; Tg - combustion temperature of fuel gas; Vc - free volume of combustion chamber cavity;

[0055] η P - is the total pressure recovery coefficient of the nozzle flow; At - throat area; C* - propellant characteristic velocity m / s; - Nozzle flow correction factor;

[0056] According to formulas (1) and (2), combined with the performance parameters of the grain and propellant, the combustion chamber working pressure P can be calculated: C , throat area At, expansion ratio ξ, working time, two-stage working time ratio, total engine impulse and two-stage impulse ratio are used as basic parameters for optimization design;

[0057] Step 2: Shell Design

[0058] Use the fourth strength theory to calculate the shell wall thickness and determine the combustion chamber shell wall thickness δ min :

[0059]

[0060] Where: —pressure fluctuation coefficient; ξ—weld strength coefficient; Di—inner diameter of the cylinder; Pmax—maximum working pressure; [σ]—allowable stress of the material, [σ]=σb / nb, this engine.

[0061] The shell wall thickness δ is obtained by design min ,At the same time, according to the shape and weight requirements, the length and detailed structure design of the combustion chamber shell are completed, and the shell weight design parameters are obtained as the basic parameters for optimization design;

[0062] Step 3: Nozzle Design

[0063] Based on the throat area At and expansion ratio calculated in step 1, the main task of nozzle design is to complete the aerodynamic design and thermal protection design of the nozzle.

[0064] Aerodynamic design is the shape parameters of the convergent section, throat and diffuser section, which makes the nozzle more efficient. The aerodynamic design of the nozzle should ensure the continuity of the profile, and adopt the form of tangent connection of the inlet arc and straight line to avoid the existence of compression waves and shock waves caused by discontinuous profile. The downstream design of the throat is mainly to select the expansion half angle of the expansion section to reduce the sudden expansion loss caused by the profile, and complete the internal profile design of the nozzle under the conditions of meeting the throat diameter, outlet diameter and external size constraints.

[0065] The throat diameter is

[0066]

[0067] The outlet diameter is

[0068] D e =Dtξ 0.5

[0069] Combined with the designed engine working pressure and working time, the design of the nozzle thermal protection structure is completed, and the design parameters such as nozzle weight, length, and expansion half angle are obtained as the basic parameters for optimized design.

[0070] Step 4: Build the optimization model

[0071] Under the constraints of performance and size, the three optimization objectives of engine mass ratio, total impulse and vehicle range are established in combination with the design and use requirements. The least squares loss function of each optimization objective is determined, which can make each independent optimization objective converge to the optimal parameter. The research object can be a large-size single-chamber dual-thrust solid engine scheme, and it is also applicable to a single-chamber large-thrust engine scheme. The optimization parameters include the engine combustion chamber pressure P obtained by the previous calculation. C , throat area At, combustion chamber shell wall thickness δ min , expansion ratio ξ, working time t, two-stage working time ratio, engine total impulse and two-stage impulse ratio, nozzle weight, length, expansion half angle and other design parameters.

[0072] The linear weighted method is used to optimize the aircraft's longest range and weight it, which can be expressed as:

[0073]

[0074] stg(x)≥0,i∈[1,M]

[0075] h j (x)=0,j∈[1,L]

[0076] Through the linear weighted algorithm, multi-objective optimization is transformed into a more achievable single-objective optimization problem model.

[0077] Step 5: Perform optimization calculations

[0078] The optimization target weights are set differently. When the range optimization is the main target, the range optimization weight can be set to 0.6, and the mass ratio and total impulse can be set to 0.2. The main optimization target is optimized. Under the constraints of engine working pressure and working time, reasonable and effective results are selected as candidate solution sets to obtain multiple candidate solutions.

[0079] Step 6: Evaluate and determine the final optimization solution

[0080] The main objective function value of the candidate target schemes is evaluated, and the engineering feasibility of the engine charge and internal ballistic scheme, shell scheme, and nozzle scheme under the optimal conditions of the main target (range) is comprehensively analyzed and evaluated to obtain the final scheme that meets the optimization expectations, meets the usage requirements, and has the ability to be engineered.

[0081] A performance optimization method for a large-sized single-chamber dual-thrust solid engine based on multi-objective optimization provided by a preferred embodiment of the present invention has been successfully experimentally verified and can meet the use requirements.

[0082] like Figure 3 After the method is used to optimize the design of a certain engine scheme, the engine mass ratio is increased by 2.8%, the corresponding engine total impulse is increased by 3.7%, and the range is increased by 4%, while the engine working time remains unchanged. The overall performance of the engine and the aircraft is improved. It can be seen from the above technical scheme that the present invention can realize the multi-objective optimization design of large-caliber solid engine schemes suitable for medium- and long-range rockets.

Claims

1. A method for optimizing the performance of a large single-chamber dual-thrust solid rocket engine based on multi-objective optimization, characterized in that The steps include: Step 1: Initial engine design: Design the engine's overall dimensions, structure, charge and internal ballistics; Complete the design of charge type, propellant performance and internal ballistic scheme according to the total engine impulse and working time requirements; Complete the combustion chamber casing design according to the engine operating pressure, size and interface requirements; Complete the design of nozzle assembly and thermal protection scheme according to the internal ballistic design results and working time requirements; Step 2: Construct an optimization model: This is used to establish an optimization algorithm under the constraints of performance, size, and weight. According to the optimization goals of the highest mass ratio, the highest total impulse, and the longest external trajectory matching range, a linear weighted algorithm is used in combination with the optimization weights. Step 3: Determine the optimization objectives and optimize them separately: obtain the single variable optimization candidate solution set of multiple objectives, the charge scheme and propellant performance selection set, throat diameter selection and expansion ratio scheme set that meet the maximum total impulse condition; Step 4: Evaluate and determine the final optimization plan: Analyze the weights of each optimization goal, and conduct comprehensive analysis and evaluation to obtain the final optimization design plan that is both objective and reasonable.

2. The method for optimizing the performance of a large single-chamber dual-thrust solid rocket engine based on multi-objective optimization according to claim 1 is characterized in that: The specific steps of step 1 are: Step 1.1: Charge and interior ballistics design; The charge design should improve the charge density while meeting the overall ballistic index requirements for the engine. On the one hand, mature charge type solutions such as tube slots, wing columns, and star holes should be selected to reduce the process complexity and development cost of the charge. On the other hand, the filling coefficient should be selected to improve the charge's adaptability to a wide temperature environment. On this basis, the charge column structure and propellant performance should be optimized to determine the charge type solution. After the charge design is completed, the internal ballistics is estimated according to the principle of parallel layer displacement and calculated according to the zero-dimensional steady-state engine internal ballistic differential equation. The calculation formula is as follows: Where: - Gas mass generation rate, in kg / s; Ka - burning rate correction factor; a - burning rate at calculated temperature; P C - Combustion chamber pressure, n-pressure index; A b - charge burning area; ρ p - propellant density; σ P - Burning rate temperature sensitivity coefficient (1 / K); T - calculation temperature; T0 - reference temperature; η c - characteristic velocity coefficient of the propellant; R g - Gas constant of fuel gas; T g - Combustion temperature of the gas; Vc - Free volume of the combustion chamber cavity; η P - is the total pressure recovery coefficient of the nozzle flow; A t - throat area; C* - characteristic velocity of propellant m / s; - Nozzle flow correction factor; According to formula (1) and formula (2), combined with the performance parameters of the grain and propellant, the combustion chamber working pressure P is calculated as C , throat area A t , expansion ratio ξ, working time, two-stage working time ratio, total engine impulse and two-stage impulse ratio are used as basic parameters for optimization design; Step 1.2: Shell design; Use the fourth strength theory to calculate the shell wall thickness and determine the combustion chamber shell wall thickness δ min : in, is the pressure fluctuation coefficient; ξ is the weld strength coefficient; D i is the inner diameter of the cylinder; P max is the maximum working pressure; [σ] is the allowable stress of the material, [σ] = σb / nb, the engine shell wall thickness δ min ,At the same time, according to the shape and weight requirements, the length and detailed structure design of the combustion chamber shell are completed, and the shell weight design parameters are obtained as the basic parameters for optimization design; Step 1.3: Nozzle Design Based on the calculated throat area A t and expansion ratio, nozzle design completes the nozzle's aerodynamic design and thermal protection design; The aerodynamic design is the shape parameters of the convergent section, throat and diffuser section, which makes the nozzle more efficient. The aerodynamic design of the nozzle ensures the continuity of the profile, and adopts the form of tangent connection between the inlet arc and the straight line to avoid the existence of compression waves and shock waves caused by the discontinuity of the profile. The downstream design of the throat is the expansion half angle selection of the expansion section to reduce the sudden expansion loss caused by the profile, and complete the internal profile design of the nozzle under the conditions of meeting the throat diameter, outlet diameter and external size constraints. The throat diameter is: Outlet diameter: D e =D t x 0.5 Combined with the designed engine operating pressure and working time, the design of the nozzle thermal protection structure is completed, and the design parameters of the nozzle weight, length and expansion half-angle are obtained as the basic parameters for optimized design.

3. The method for optimizing the performance of a large single-chamber dual-thrust solid rocket engine based on multi-objective optimization according to claim 1 is characterized in that: The specific steps of constructing the optimization model in step 2 are: Under the constraints of performance and size, three optimization objectives, namely engine mass ratio, total impulse and aircraft range, are established in combination with design and use requirements, and the least squares loss function of each optimization objective is determined so that each independent optimization objective converges to the optimal parameter. The research object is a large-size single-chamber dual-thrust solid engine scheme, or a single-chamber large-thrust engine scheme. The optimization parameters include the engine combustion chamber pressure P obtained by the previous calculation. C , throat area A t 、Combustion chamber shell wall thickness δ min , expansion ratio ξ, working time t, two-stage working time ratio, total engine impulse and two-stage impulse ratio, nozzle weight, nozzle length and expansion half angle; The linear weighted method is used to optimize the aircraft's longest range and weight it, which can be expressed as: stg(x)≥0,i∈[1,M] h j (x)=0,j∈[1,L] Through the linear weighted algorithm, multi-objective optimization is transformed into a more achievable single-objective optimization problem model.

4. The method for optimizing the performance of a large single-chamber dual-thrust solid rocket engine based on multi-objective optimization according to claim 1 is characterized in that: The specific steps of step 3 are: The optimization target weights are set differently. When the range optimization is the main target, the range optimization weight is set to 0.6, the mass ratio and total impulse are set to 0.2, and the main optimization target is optimized. Under the constraints of engine working pressure and working time, the best result is selected as the candidate solution set to obtain multiple candidate solutions.

5. The method for optimizing the performance of a large single-chamber dual-thrust solid rocket engine based on multi-objective optimization according to claim 1 is characterized in that: In step 4, the steps of evaluating and determining the final optimization solution are: Evaluate the main objective function value of the candidate target solutions obtained, and comprehensively analyze and evaluate the engineering feasibility of the engine charge and internal ballistics solution, shell solution, and nozzle solution under the optimal condition of the main target range, so as to obtain the final solution that meets the optimization expectations, meets the usage requirements, and has the ability to be engineered.

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