A combined propulsion device modal conversion scheme design method

Abstract

The application discloses a combined propulsion device mode conversion scheme design method, comprising the following steps: obtaining a Mach number interval of a combined propulsion device when the combined propulsion device is converted from a first mode to a second mode, and selecting a plurality of representative flight state points; for each representative flight state point, respectively performing discrete scanning on a primary propulsion unit mass flow range and a secondary fuel fuel-air ratio range, to form a plurality of candidate working condition combinations; for each candidate working condition combination, calculating flow and combustion characteristics in a thrust chamber, and screening a target working condition combination which can stably combust and meet a thrust demand; obtaining an optimal target working condition combination under each representative flight state point based on enumeration search, and generating a continuous and implementable mode conversion scheduling path. The application is applied to the combined propulsion field, and can effectively realize efficient, stable and safe operation of a mode conversion process.

Description

A design method for mode conversion scheme of a combined propulsion device Technical Field

[0001] This invention relates to the field of combined propulsion technology, specifically a design method for a mode conversion scheme of a combined propulsion device. Background Technology

[0002] Currently, engineering designs for the transition from the first to the second mode of a combined propulsion system often employ empirical or trial-and-error methods. Specifically, the general approach is as follows: First, based on the predetermined flight trajectory and using engineering experience, empirical formulas, or simplified aerodynamic / propulsion estimation methods, a single typical flight state point is selected, such as the dynamic pressure or altitude point at Ma≈3. The total thrust required by the combined propulsion system in this state, as well as the thrust shares borne by the primary and secondary propulsion units, are estimated. Based on this, through simple performance estimation or single-point CFD calculations, the mass flow rate of the primary propulsion unit, the secondary fuel-air ratio, and the approximate thrust chamber pressure distribution corresponding to this state point are derived.

[0003] In some design schemes, the first and second modes are even optimized independently. That is, the mass flow rate of the primary propulsion unit and the secondary fuel air-fuel ratio are optimized separately in the first mode range, and the secondary propulsion scheme is optimized separately in the second mode range. Finally, a switching point is selected based on engineering experience to simply connect the two modes, and the entire transition range is rarely designed as a whole and the constraints are checked.

[0004] While existing traditional technologies have some feasibility in engineering, they still have significant limitations:

[0005] First, it is usually designed based on a single typical flight state point such as Ma≈3, which fails to reflect the parameter evolution law of continuous change with Mach number and dynamic pressure in the transition interval between the first mode and the second mode, resulting in insufficient robustness of the scheme to flight trajectory deviation and environmental changes.

[0006] Secondly, key parameters such as primary propulsion unit mass flow rate and secondary fuel air-fuel ratio are mostly obtained from experience or simplified estimation, lacking system optimization under multiple operating conditions and constraints, making it difficult to simultaneously consider indicators such as thrust demand, specific impulse performance, combustion stability and thrust chamber heat load.

[0007] Furthermore, the first and second modes are often optimized separately, with only a simple connection at a single empirical "switching point." There is a lack of overall inspection and evaluation of engineering constraints such as thrust continuity, smoothness of secondary fuel air-fuel ratio changes, and propellant supply and valve adjustment capabilities throughout the entire transition range. This can easily lead to risks such as sudden thrust changes, parameter jumps, or local operating condition overruns during mode transitions, making it difficult to fully guarantee engineering applicability and safety margins. Summary of the Invention

[0008] To address the shortcomings of the existing technology, this invention provides a design method for mode conversion schemes of combined propulsion devices. This method comprehensively considers flight Mach number, thrust requirements, secondary fuel fuel-air ratio constraints, and supply system capabilities to determine the optimal matching relationship between the primary propulsion unit mass flow rate and the secondary fuel fuel-air ratio, and generates an engineering-feasible mode conversion path scheme, thereby effectively achieving efficient, stable, and safe operation of the mode conversion process.

[0009] To achieve the above objectives, the present invention provides a design method for a mode conversion scheme of a combined propulsion device, comprising the following steps:

[0010] Step 1: Obtain the Mach number range of the combined propulsion device from the first mode to the second mode, and select several representative flight state points within the Mach number range;

[0011] Step 2: For each representative flight state point, perform discrete scanning on the range of primary propulsion unit mass flow rate and the range of secondary fuel air-fuel ratio to obtain several candidate primary propulsion unit mass flow rates and several candidate secondary fuel air-fuel ratios. Then, combine the representative flight state point with each candidate primary propulsion unit mass flow rate and each candidate secondary fuel air-fuel ratio to form several candidate operating condition combinations.

[0012] Step 3: For each candidate working condition combination, calculate the flow and combustion characteristics in the thrust chamber of the combined propulsion device using experiments or numerical simulations, and screen out all target working condition combinations that can achieve stable combustion and meet the thrust requirements.

[0013] Step 4: Based on the flight mission objective, obtain the optimal combination of target conditions for each representative flight state point through enumeration search.

[0014] Step 5: Based on the optimal target condition combination of each representative flight state point, generate a continuously implementable mode conversion scheduling path.

[0015] Compared with the prior art, the present invention has the following beneficial technical effects:

[0016] 1. This invention selects multiple representative flight state points within the mode transition interval and scans the primary propulsion unit mass flow rate and secondary fuel fuel-air ratio to construct a complete multi-dimensional matching database of flight Mach number-primary propulsion unit mass flow rate-secondary fuel fuel-air ratio. This improves upon the traditional single-point design to a unified design across multiple state intervals, effectively reflecting the parameter evolution law within the entire transition interval and exhibiting better robustness to flight trajectory disturbances and environmental changes.

[0017] 2. In the implementation phase, this invention can uniformly superimpose multiple engineering constraints such as thrust requirements, upper and lower limits of secondary fuel air-fuel ratio, thermal load limits, propellant supply, and valve regulation capabilities, systematically screen candidate operating condition combinations, and flexibly determine optimization objectives according to different mission requirements. This realizes the transformation from simply pursuing thrust maximization to optimizing comprehensive performance, making the design results better adaptable to diverse mission requirements. Compared with the existing technology that simply sets switching points based on engineering experience and lacks complete constraint checks, this invention can better ensure thrust continuity, combustion stability, and thermal safety during mode transition, and improve the feasibility and safety margin of the scheme.

[0018] 3. During the implementation phase, this invention can generate a smooth scheduling path that falls within the feasible area based on the optimal operating conditions and the capabilities of the supply system and control system. This ensures that the flow rate of the primary propulsion unit and the air-fuel ratio of the secondary fuel change continuously with the Mach number, significantly reducing the risk of thrust mutation and parameter jump. It also facilitates the implementation of the control law, reduces the workload of repeated ground tests and scheme iterations, and improves the efficiency and reliability of mode conversion scheme design. Attached Figure Description

[0019] 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.

[0020] Figure 1 is a flowchart of the design method for the mode conversion scheme of the combined propulsion device in an embodiment of the present invention;

[0021] Figure 2 is a schematic diagram of the feasible adjustment range of the primary propulsion unit mass flow rate and the secondary fuel air-fuel ratio under the Ma=3.25 condition in an embodiment of the present invention.

[0022] 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

[0023] 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.

[0024] Furthermore, the technical solutions of the various embodiments of the present invention can be combined with each other, but only if they are based on the ability of those skilled in the art to implement them. When 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.

[0025] This embodiment discloses a design method for a mode conversion scheme of a combined propulsion device. The combined propulsion device consists of a primary propulsion unit and a secondary propulsion unit. The primary propulsion unit mainly generates thrust by reacting and burning the fuel carried by the aircraft itself with an oxidizer. The secondary propulsion unit generates thrust by reacting and burning the fuel carried by the aircraft itself with the airflow introduced through the air intake. The combined propulsion device has a first mode and a second mode. The first mode corresponds to the startup phase of the combined propulsion device, at which time the airflow cannot provide sufficient pressure for the propulsion device; that is, in the first mode, thrust is provided only by the primary propulsion unit. The second mode corresponds to the normal flight phase of the combined propulsion device, at which time the airflow can provide sufficient pressure for the propulsion device. Thrust is generated by injecting secondary fuel into the secondary propulsion unit and reacting and burning with oxygen in the airflow; that is, in the second mode, thrust is provided only by the secondary propulsion unit.

[0026] When the combined propulsion system transitions from the first mode to the second mode, there is a switching process where the primary propulsion unit shuts down and the secondary propulsion unit starts. This embodiment proposes a design method for the mode transition scheme of the combined propulsion system. Given the flight mission and combined propulsion system parameters, this method can determine the matching relationship between the mass flow rate of the primary propulsion unit and the fuel-air ratio of the secondary propulsion unit, and form a continuous mode transition scheduling path that satisfies engineering constraints. Referring to Figure 1, the specific steps of the combined propulsion system mode transition scheme design method in this embodiment are as follows:

[0027] Step 1: Obtain the Mach number range of the combined propulsion device from the first mode to the second mode, and select several representative flight state points within the Mach number range. That is, the entire Mach number range of the mode transition is discretized and used for subsequent parameter scanning and optimization studies.

[0028] Step 2: For each representative flight state point, perform discrete scanning of the primary propulsion unit mass flow rate range and the secondary fuel air-fuel ratio range to obtain several candidate primary propulsion unit mass flow rates and several candidate secondary fuel air-fuel ratios. Then, combine the representative flight state point with each candidate primary propulsion unit mass flow rate and each candidate secondary fuel air-fuel ratio to form several candidate operating condition combinations.

[0029] Step 3: For each candidate working condition combination, use experiments or numerical simulations to calculate the flow and combustion characteristics in the thrust chamber of the combined propulsion device, and screen out all target working condition combinations that can achieve stable combustion and meet the thrust requirements.

[0030] Step 4: Based on the flight mission objectives, obtain the optimal combination of target conditions for each representative flight state point through enumeration search.

[0031] Step 5: Based on the optimal combination of target operating conditions for each representative flight state point, generate a continuous and feasible mode conversion scheduling path.

[0032] In this embodiment, the thrust gain coefficient is used as a criterion to determine the upper boundary of the Mach number interval for the transition from the first mode to the second mode. The thrust gain coefficient is defined as the ratio of the measured total thrust of the combined propulsion device to the theoretical thrust of a single propulsion unit operating alone. In specific implementation, the process of determining the upper boundary of the Mach number interval is as follows:

[0033] When the thrust gain coefficient decreases below the first threshold as the flight Mach number increases, it is considered that the thrust gain generated by the first mode ejection is insufficient to compensate for the propellant consumption cost caused by the high flow rate operation of the propulsion unit. The first mode no longer has an advantage. Therefore, when the thrust gain coefficient decreases below the first threshold as the flight Mach number increases, the flight Mach number at this point is the upper boundary of the Mach number range. The first threshold can be set to 1.3~1.5.

[0034] In this embodiment, the lower boundary of the Mach number interval for the transition from the first mode to the second mode is determined by combining the activation criterion of the upper air intake channel of the secondary propulsion unit and the combustion stability criterion of the upper thrust chamber of the secondary propulsion unit. The specific implementation process is as follows:

[0035] When the average Mach number at the inlet section of the thrust chamber in the secondary propulsion unit is consistently below 0.5, the amplitude of the pressure pulsation on the wall of the isolation section in the secondary propulsion unit is less than 5% of its average pressure, and the amplitude of the pressure pulsation on the wall of the thrust chamber in the secondary propulsion unit is less than 10% of its average pressure, the flight Mach number at this time is the lower boundary of the Mach number range.

[0036] In the specific implementation process of selecting representative flight state points, five representative flight state points were selected based on the evolution law of the working characteristics of the combined propulsion device in the Mach number range. The first representative flight state point was set as the lower boundary of the Mach number range, representing the typical characteristics of being completely in the first mode. At this time, the primary propulsion unit provides all or most of the thrust, and the secondary propulsion unit has not yet started or is in standby mode. The Mach number of the second representative flight state point was set as the lower boundary value of the Mach number range plus 0.05~0.15, representing the initial characteristics of the transition from the first mode to the mixed mode (i.e., the primary propulsion unit and the secondary propulsion unit operate simultaneously). At this time, the contribution of the primary propulsion unit is still relatively high, and the role of the secondary propulsion unit begins to appear. The first representative flight state point is set at the median of the Mach number range, representing a transitional characteristic where the primary and secondary propulsion units are relatively balanced, and is a key node where the two working mechanisms alternately dominate. The second representative flight state point is set at the upper boundary of the Mach number range minus 0.05 to 0.15, representing the terminal characteristic of the second mode being established. At this time, the secondary propulsion unit has taken the lead, and the role of the primary propulsion unit gradually decreases to an auxiliary propulsion function. The third representative flight state point is set at the upper boundary of the Mach number range, representing the stable characteristic of fully entering the second mode. At this time, the secondary propulsion unit has undertaken the main propulsion task, and the primary propulsion unit is in a closed state or in a standby state. For example, when the Mach number range is Mach 3 to Mach 3.5, the selected representative flight state points are Mach 3, Mach 3.1, Mach 3.25, Mach 3.4, and Mach 3.5, respectively.

[0037] In the specific implementation of step 2, for different representative flight state points, the discretization criteria for the primary propulsion unit mass flow rate range and the secondary fuel fuel-air ratio range are the same. For example, if the primary propulsion unit mass flow rate range is 1 kg / s to 7 kg / s and the secondary fuel fuel-air ratio range is 0.2 to 1.4, for the five representative flight state points Mach 3, Mach 3.1, Mach 3.25, Mach 3.4, and Mach 3.5, the primary propulsion unit mass flow rate range is discretized into seven candidate primary propulsion unit mass flow rates of 1 kg / s, 2 kg / s, 3 kg / s, 4 kg / s, 5 kg / s, 6 kg / s, and 7 kg / s. At the same time, the secondary fuel fuel-air ratio range is discretized into seven candidate secondary fuel fuel-air ratios of 0.2, 0.4, 0.6, 0.8, 1, 1.2, and 1.4. Finally, 245 candidate operating condition combinations containing flight state (Ma), primary propulsion unit mass flow rate (kg / s), and secondary fuel fuel-air ratio can be formed.

[0038] It is worth noting that in specific applications, it is not limited to selecting the above five representative flight state points. More other flight state points can also be selected based on the above five representative flight state points. For example, based on the above five representative flight state points, the median between some or all of the two adjacent representative flight state points can also be used as representative flight state points, that is, one or more of Mach 3.05, Mach 3.175, Mach 3.325, and Mach 3.45 can also be used as representative flight state points.

[0039] This embodiment employs the same discretization standard for the primary propulsion unit mass flow rate range and the secondary fuel-air ratio range for different representative flight state points. This not only ensures that the data at each representative flight state point have the same parameter dimensions and sampling density, facilitating the establishment of a structured multidimensional matching database and simplifying subsequent data storage, retrieval, and batch processing, but also promotes performance comparison analysis across flight state points. It allows for intuitive observation of the evolution of combined propulsion device performance with flight Mach number under the same primary propulsion unit mass flow rate conditions, making it easier to identify performance change trends during mode transitions. Furthermore, the unified discretization standard significantly simplifies interpolation calculations in the subsequent scheduling path generation step. Since each representative flight state point has the same flow rate sampling grid, point-to-point interpolation can be performed directly without complex non-uniform grid resampling or scattered point interpolation processing. This improves the automation and repeatability of the overall calculation process, enabling batch numerical simulation task scheduling, significantly improving computational efficiency and reducing the need for manual intervention.

[0040] In this embodiment, the specific implementation process of step 3 is as follows:

[0041] First, for each candidate operating condition combination, the flow and combustion characteristics in the thrust chamber of the combined propulsion device are calculated using experiments or numerical simulations to obtain the performance indicators of the combined propulsion device under each candidate operating condition combination, including thrust, specific impulse, total pressure recovery of the thrust chamber, wall heat load, etc.

[0042] Then, all candidate operating condition combinations are screened for the first time. Specifically, during the experimental or numerical simulation of each candidate operating condition combination, if the pressure oscillation amplitude of the thrust chamber in the combined propulsion device does not exceed 10% of the average pressure, the candidate operating condition combination is determined to be able to burn stably and is defined as a stable combustion operating condition combination. For example, as shown in Figure 2, the area covered by the horizontal and vertical axes is all candidate operating condition combinations under the representative flight state point corresponding to Ma=3.25. Stage 1 is the feasible region after being clipped by the constraint condition of stable combustion in the primary propulsion unit mass flow rate-secondary fuel air-fuel ratio plane. This region constitutes the feasible adjustment framework in the first mode and second mode transition interval, which corresponds to the stable combustion operating condition combination.

[0043] Finally, based on the target thrust or allowable thrust range and engineering constraints corresponding to each representative flight state point given by the flight mission requirements, each stable combustion condition combination is checked for constraints one by one. Stable combustion condition combinations that meet the thrust requirements and engineering constraints are selected and output as the target condition combination. Among them, the engineering constraints include the maximum and minimum values ​​of the primary propulsion unit mass flow rate, and the upper and lower limits of the secondary fuel air-fuel ratio, etc. For example, stage 2 shown in Figure 2 is the target region after being clipped by the target thrust or allowable thrust range and engineering constraints.

[0044] In the specific implementation of step 4, single-objective or multi-objective optimization functions can be set according to different engineering requirements. For example, for flight missions emphasizing rapid maneuverability and climb capability, the objective can be set to maximize thrust; for long-range cruise missions emphasizing range capability, the objective can be set to maximize specific impulse (i.e., minimize propellant consumption) while meeting the lower thrust limit; and for reusable aircraft operating for extended periods, the objective can be set to minimize thrust chamber heat load and pressure fluctuation amplitude while ensuring that thrust and specific impulse meet the lower limits. For each representative flight state point, an enumeration search is used among all corresponding target condition combinations to optimize the preset objective function, obtaining the operating point with the optimal performance index under the constraints, i.e., the corresponding optimal target condition combination.

[0045] In the specific implementation of step 5, the mode transition scheduling path needs to be generated based on the optimal target operating condition combination of each representative flight state point and combined with the engineering characteristics of the propellant supply and control system. Specifically, the actual capability constraints of the supply system can be quantified into the following typical indicators: In the propellant supply system of the primary propulsion unit, the valve opening change rate should be controlled within 15% / s of the full stroke, and the response delay time is about 50ms~150ms; In the secondary fuel supply branch, the secondary fuel air-fuel ratio change rate should be controlled within 0.05 / s, and the regulating valve response time is usually 80ms~200ms; Considering that the typical available time window for mode transition of the aircraft in the Mach number range of 3.0~3.5 is about 5s~15s, the parameter adjustment path between each representative flight state point needs to complete a smooth transition under this time constraint.

[0046] Considering engineering factors such as the upper limit of the flow regulation rate of the primary propulsion unit and the secondary fuel channel, the valve opening resolution and response time, and the linkage constraints between the primary propulsion unit and the secondary fuel flow, this embodiment uses the discrete optimal operating points (i.e., the optimal target operating condition combination) under each representative flight state point as a reference. It employs polynomial fitting, spline curves, or piecewise functions to construct scheduling curves for the primary propulsion unit mass flow rate as a function of Mach number (or time) and for the secondary fuel fuel-air ratio as a function of Mach number. During the fitting process, the scheduling curves are required to remain within a feasible adjustment framework within the transition range between the first and second modes, avoiding contact with constraints such as thrust, secondary fuel fuel-air ratio, heat load, and supply capacity. When determining the transition path, the rate of change of the primary propulsion unit mass flow rate between adjacent flight state points must not exceed the maximum regulation speed of the supply system; otherwise, it will lead to supply pressure instability or abnormal injector operation. Simultaneously, the flow rate change trajectory should be as smooth as possible, avoiding abrupt changes or oscillations to reduce disturbances to combustion stability. When the fitted curve poses a risk of exceeding limits or experiences excessively rapid flow changes in local areas, making it unsuitable for engineering implementation, it can be appropriately modified to deviate from the theoretical optimum by a certain margin, thereby achieving better feasibility and control smoothness. Ultimately, a continuous scheduling path that gradually evolves from the first mode of operation to the second mode of operation can be obtained.

[0047] Taking cubic spline curve fitting as an example, there are n representative flight state points with Mach numbers Ma1, Ma2, ..., Ma... n The corresponding optimal first-stage propulsion unit mass flow rates are respectively 1. 2、···、 n The fitting process for the mass flow scheduling curve of a single propulsion unit in a continuous scheduling path is as follows:

[0048] First, a cubic polynomial is defined between each pair of adjacent data points. That is, the entire mass flow scheduling curve of a single propulsion unit is composed of n-1 segments of cubic polynomials. Segment polynomial (connecting the first) The point and the first The general form of (a number of points) is:

[0049] ;

[0050] in, , , , There are 4(n-1) unknown coefficients to be determined.

[0051] Then, a set of conditional equations is established to determine the coefficients. The conditions include: interpolation conditions, i.e., the curve must pass through all data points, and the function values ​​of each polynomial at its starting and ending points must be equal to the corresponding data values, providing 2(n-1) equations; continuity conditions, i.e., the function values, first derivative values, and second derivative values ​​of two adjacent polynomials at common nodes must be equal, providing 3(n-2) equations; boundary conditions, used to determine the behavior of the curve at both ends. Commonly used boundary conditions include natural boundary conditions (the second derivative at the endpoints is zero) or clamping boundary conditions (specifying the first derivative value at the endpoints), providing 2 equations; the above conditions provide a total of 4(n-1) equations, determining all 4(n-1) unknown coefficients;

[0052] Subsequently, the above conditions are rearranged into a system of linear equations with unknown coefficients. This system of equations has a banded sparse structure and can be solved using efficient algorithms such as the chasing method. For example, the constant terms of each segment can be determined first using interpolation conditions. The function value is equal to the corresponding data point; then a tridiagonal system of equations is established about the second derivative values ​​of each node, and the second derivative values ​​of each node are obtained by solving the system of equations; finally, the remaining coefficients of each segment are deduced based on the second derivative values ​​and interpolation conditions.

[0053] After the solution is obtained, a feasible region constraint check is performed, which checks whether each point on the curve is within the defined feasible adjustment framework. Specifically, this involves densely sampling several check points within the Mach number range, calculating the mass flow rate of each propulsion unit corresponding to each check point, and determining whether these values ​​fall within the feasible flow rate range corresponding to that Mach number. If the curve is found to exceed the feasible region boundary in certain areas, the curve needs to be adjusted. Adjustment methods include: adding constraint points near the out-of-limit region to force the curve to pass through points within the feasible region, and then re-fitting the spline; or using a constrained least squares fitting method, directly applying feasible region boundary constraints during the fitting process.

[0054] Finally, a rate of change constraint test and smoothing process are performed. This involves checking whether the rate of change of the fitted curve meets the engineering implementation constraints. For example, the first derivative of the curve is calculated to check whether the rate of change of the mass flow rate of the primary propulsion unit with Mach number exceeds the maximum allowable adjustment rate of the supply system (the Mach number rate of change needs to be converted to a time rate of change based on the flight trajectory). If the rate of change exceeds the limit, the following methods can be used: increase the Mach number interval between adjacent data points to reduce the local slope of the curve; or perform low-pass filtering on the curve in the excess region to smooth out overly drastic changes; or replace the spline curve in that region with a piecewise linear function to control the rate of change more directly.

[0055] The fitting process for the secondary fuel air-fuel ratio scheduling curve is entirely similar to that for the primary propulsion unit mass flow rate scheduling curve. Using the same spline fitting method, the optimal secondary fuel air-fuel ratio data points are fitted as a continuous curve about the Mach number, and corresponding constraint checks and adjustments are performed. Ultimately, the primary propulsion unit mass flow rate scheduling curve and the secondary fuel air-fuel ratio scheduling curve together constitute a complete mode transition scheduling path.

[0056] In summary, the mode conversion scheme design method of the combined propulsion device in this embodiment, under the constraints of a given geometric thrust chamber and a predetermined flight mission, through parameter scanning, constraint screening, optimal operating condition search, and path smoothing fitting of representative flight state points, can provide the optimal matching relationship and continuous scheduling path between the mass flow rate of the primary propulsion unit and the fuel-air ratio of the secondary fuel, while ensuring various engineering constraints such as thrust requirements, fuel-air ratio constraints, thrust chamber thermal safety, and propellant supply capacity. This achieves efficient, stable, and safe operation of the combined propulsion device during mode conversion, overcoming the shortcomings of traditional design methods that rely on single-point experience and simple switching.

[0057] The above description is only a preferred embodiment of the present invention and does not limit the scope of protection of the present invention. All equivalent structural transformations made under the inventive concept of the present invention using the contents of the present invention specification and drawings, or direct / indirect applications in other related technical fields, are included within the scope of protection of the present invention.

Claims

1. A design method for a mode conversion scheme of a combined propulsion device, characterized in that, The process includes the following steps: Step 1, obtaining the Mach number range for the combined propulsion device's transition from the first mode to the second mode, and selecting several representative flight state points within the Mach number range. The selected representative flight state points include: the first representative flight state point, set as the lower boundary of the Mach number range, representing a typical characteristic of being completely in the first mode, where the primary propulsion unit provides all or most of the thrust, and the secondary propulsion unit has not yet started or is in standby mode; the second representative flight state point, whose Mach number is set to the lower boundary value of the Mach number range plus 0.05~0. 0.15 represents the initial characteristic of the transition from the first mode to the mixed mode. At this time, the contribution of the primary propulsion unit is still relatively high, and the role of the secondary propulsion unit begins to appear but has not yet become dominant. The third representative flight state point is set to the median of the Mach number range, representing the transition characteristic of a relatively balanced role between the primary and secondary propulsion units. The fourth representative flight state point has its Mach number set to the upper boundary value of the Mach number range minus 0.05 to 0.15, representing the terminal characteristic of the second mode being established. At this time, the secondary propulsion unit has become dominant, and the primary propulsion unit plays a less dominant role. The process is as follows: The propulsion function is gradually reduced to auxiliary propulsion. The fifth representative flight state point is set as the upper boundary of the Mach number range, representing the stable characteristics of fully entering the second mode. At this time, the secondary propulsion unit has undertaken the main propulsion task, and the primary propulsion unit is in a closed or standby state. Step 2: For each representative flight state point, the mass flow rate range of the primary propulsion unit and the secondary fuel-air ratio range are discretely scanned to obtain several candidate primary propulsion unit mass flow rates and several candidate secondary fuel-air ratios. The representative flight state point is then combined with each candidate primary propulsion unit mass flow rate and each candidate secondary fuel-air ratio to form several candidate operating condition combinations. Step 3: For each candidate operating condition combination, the flow and combustion characteristics in the thrust chamber of the combined propulsion device are calculated using experiments or numerical simulations, and all target operating condition combinations that can achieve stable combustion and meet thrust requirements are selected. Step 4: Based on the flight mission objective, the optimal target operating condition combination under each representative flight state point is obtained through enumeration search. Step 5: Based on the optimal target operating condition combination of each representative flight state point, a continuously implementable mode transition scheduling path is generated.

2. The design method for the mode conversion scheme of the combined propulsion device according to claim 1, characterized in that, In step 1, the thrust gain coefficient is used as a criterion to determine the upper boundary of the Mach number interval. The thrust gain coefficient is the ratio of the measured total thrust of the combined propulsion device to the theoretical thrust when the primary propulsion unit works alone. The process of determining the upper boundary of the Mach number interval is as follows: when the thrust gain coefficient decreases to below the first threshold as the flight Mach number increases, the flight Mach number at this time is the upper boundary of the Mach number interval.

3. The design method for the mode conversion scheme of the combined propulsion device according to claim 1, characterized in that, In step 1, the lower boundary of the Mach number range is determined by combining the start-up criteria of the upper air intake channel of the secondary propulsion unit and the combustion stability criteria of the upper thrust chamber of the secondary propulsion unit. Specifically, when the average Mach number of the inlet section of the upper thrust chamber of the secondary propulsion unit is continuously lower than 0.5, the pressure pulsation amplitude of the wall of the upper isolation section of the secondary propulsion unit is less than 5% of its average pressure, and the pressure pulsation amplitude of the wall of the upper thrust chamber of the secondary propulsion unit is less than 10% of its average pressure, the flight Mach number at this time is the lower boundary of the Mach number range.

4. The design method for the mode conversion scheme of the combined propulsion device according to claim 1, 2, or 3, characterized in that, In step 2, the same discrete standard is used for the range of mass flow rate of the primary propulsion unit and the range of air-fuel ratio of the secondary fuel for different representative flight state points.

5. The design method for the mode conversion scheme of the combined propulsion device according to claim 1, 2, or 3, characterized in that, Step 3 specifically includes: for each candidate operating condition combination, calculating the flow and combustion characteristics in the thrust chamber of the combined propulsion device using experiments or numerical simulations to obtain the performance indicators of the combined propulsion device under each candidate operating condition combination; during the experiment or numerical simulation of each candidate operating condition combination, if the pressure oscillation amplitude in the thrust chamber of the combined propulsion device does not exceed 10% of the average pressure, then the candidate operating condition combination is determined to be able to burn stably and is defined as a stable combustion operating condition combination; based on the target thrust or allowable thrust range and engineering constraints corresponding to each representative flight state point given by the flight mission requirements, each stable combustion operating condition combination is checked for constraints one by one, and stable combustion operating condition combinations that meet the thrust requirements and engineering constraints are selected and output as the target operating condition combination.

6. The design method for the mode conversion scheme of the combined propulsion device according to claim 1, 2, or 3, characterized in that, In step 5, the mode conversion scheduling path includes the primary propulsion unit mass flow rate scheduling curve and the secondary fuel air-fuel ratio scheduling curve; the primary propulsion unit mass flow rate scheduling curve and the secondary fuel air-fuel ratio scheduling curve are obtained by polynomial fitting, spline curve fitting or piecewise function fitting.

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