A dual-modal axisymmetric combined nozzle, its design method, and related equipment

Abstract

This invention discloses a dual-mode axisymmetric combined nozzle, its design method, and related equipment. The method involves removing the approximately straight portion of the initial expansion section profile and using the remaining portion as the expansion section profile of the rocket-mode Laval nozzle. A straight line perpendicular to the axis of the reference nozzle is drawn outwards from the end of the initial expansion section profile. A first arc is drawn upstream of the exhaust gas flow, centered on a point on this line. This first arc serves as the contraction section profile of the intake-mode Laval nozzle. The first arc must satisfy the following condition: the first arc and the intake-mode... The combustion chamber outlet is connected, and the expansion section profile of the rocket-mode Laval nozzle, considering its thickness, matches the throat area formed by the first arc to the exhaust gas flow rate at the intake mode combustion chamber outlet. Taking a point on the straight line as the center and the end of the first arc furthest from the intake mode combustion chamber outlet as the starting point, a second arc is drawn downstream of the exhaust gas flow, tangent to the wave-damping section profile. The portion of the wave-damping section profile before the tangent point is removed, and the remaining portion of the second arc and the wave-damping section profile is used as the expansion section profile of the intake mode Laval nozzle. This invention can meet stringent dimensional constraints, achieve high dual-channel coupling, high surface continuity, and effectively organize the flow field structure.

Description

Technical Field

[0001] This invention belongs to the field of tail nozzle design technology, specifically relating to a dual-modal axisymmetric combined nozzle, its design method, and related equipment. Background Technology

[0002] The nozzle is a crucial component of aerospace propulsion systems, and its aerodynamic performance and envelope dimensions significantly impact the design and performance of both the engine and the spacecraft. Using high-specific-impulse air-breathing engines at sea level and low altitudes, while employing high-thrust-to-weight-ratio rocket engines at high altitudes, is a feasible combined propulsion scheme suitable for the first stage of two-stage-to-orbit spacecraft. The SABRE engine (Synergetic Air-Breathing Rocket Engine) is a typical example of this combined approach. Its nozzle features a dual-throat design, achieving flexible switching between modes through the forward and backward movement of the bell-shaped nozzle's rear section. However, it retains the complete bell-shaped nozzle profile, leaving room for improvement in weight reduction.

[0003] Currently, there is no systematic approach to the profile design of dual-mode nozzles. There are two main approaches: one is to directly divide the ideal nozzle, but this method does not incorporate the profile of the outer bypass duct's convergence section into the design process and has significant mass; the other is to modify the profile based on the ideal nozzle, according to the flow rate and operating point area ratio. The disadvantages of this approach are low coupling in the design process, poor continuity between the two channels, and a large outlet size. Experiments have shown that the nozzle profiles obtained by existing design methods will cause flow separation in the inner thrust chamber when operating together, and the outer bypass duct nozzle will cause significant interference to the internal combustion gas in a vacuum environment.

[0004] For dual-mode engines with a wide flight envelope, it is necessary to develop a dual-mode combined nozzle profile design method that can meet stringent size constraints, have a high degree of dual-channel coupling, high profile continuity, and effectively organize the flow field structure. Summary of the Invention

[0005] To address the problems existing in the prior art, the present invention provides a dual-modal axisymmetric combined nozzle, design method and related equipment, which can meet strict size constraints, high degree of dual-channel coupling, high surface continuity and effectively organize the flow field structure.

[0006] To solve the above-mentioned technical problems, the present invention is achieved through the following technical solution:

[0007] According to a first aspect of the present invention, a design method for a dual-mode axisymmetric combined nozzle is provided, the combined nozzle comprising a rocket-mode Laval nozzle and an intake-mode Laval nozzle sleeved outside the rocket-mode Laval nozzle, the design method comprising:

[0008] The initial expansion section profile and the wave-damping section profile of the reference nozzle are obtained. The initial expansion section profile is determined based on the analytical design method and the characteristic line method, and the wave-damping section profile is determined based on the maximum thrust theory. The reference nozzle refers to a single-channel Laval nozzle matched only when considering rocket modes.

[0009] Remove the approximately straight portion of the initial expansion section profile and use the remaining portion of the initial expansion section profile as the expansion section profile of the rocket mode Laval nozzle;

[0010] Draw a straight line perpendicular to the reference nozzle axis from the end of the initial expansion section profile. With a point on the straight line as the center, draw a first arc upstream of the gas flow. Use the first arc as the contraction section profile of the intake mode Laval nozzle. The first arc must satisfy the following: the first arc connects to the outlet of the intake mode combustion chamber, and the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc, when considering the thickness, matches the gas flow rate at the outlet of the intake mode combustion chamber.

[0011] Taking the point on the straight line as the center and the end of the first arc away from the outlet of the intake mode combustion chamber as the starting point, draw a second arc tangent to the wave-damping section profile downstream of the gas flow. Remove the part of the wave-damping section profile before the tangent point, and use the remaining part of the second arc and the wave-damping section profile as the expansion section profile of the intake mode Laval nozzle.

[0012] In one possible implementation of the first aspect, the process of determining the initial expansion section profile of the reference nozzle specifically includes:

[0013] The core point Mach number of the reference nozzle is determined by the method of characteristics. The coordinates and airflow angles at the end of the initial expansion section profile are iterated until the core point Mach number of the reference nozzle is close to the exit Mach number of the reference nozzle under the predetermined ideal state. Then the initial expansion section profile of the reference nozzle is determined.

[0014] In one possible implementation of the first aspect, the method for determining the exit Mach number of the reference nozzle under ideal conditions includes:

[0015] Based on the geometric constraints of the dual-mode axisymmetric combined nozzle, determine the inlet and outlet dimensions of the dual-mode axisymmetric combined nozzle;

[0016] The throat area of ​​the reference nozzle is determined based on the exhaust gas flow rate at the rocket's modal combustion chamber outlet.

[0017] Based on the throat area of ​​the reference nozzle and the exit size of the dual-mode axisymmetric combined nozzle, the exit Mach number of the reference nozzle under ideal conditions is determined using one-dimensional flow theory.

[0018] In one possible implementation of the first aspect, the process of determining the profile of the damping section of the reference nozzle specifically includes:

[0019] The location of the key point of the reference nozzle is determined on the last rightward characteristic line emanating from the end of the initial expansion section profile. Combining the maximum thrust theory, the location of the key point is iterated using the bisection method to make the exit size of the reference nozzle conform to the exit size constraint of the dual-mode axisymmetric combined nozzle. The wave-damping section profile of the reference nozzle is finally determined using the unit process of the change point method and the flow balance relationship.

[0020] In one possible implementation of the first aspect, after obtaining the initial expansion section profile of the reference nozzle, the method further includes:

[0021] The contraction section profile of the reference nozzle is determined by combining the exit dimensions of the rocket mode combustion chamber. One end of the contraction section profile of the reference nozzle is connected to the exit of the rocket mode combustion chamber, and the other end smoothly transitions to the initial expansion section profile. The contraction section profile of the reference nozzle is used as the contraction section profile of the rocket mode Laval nozzle.

[0022] In one possible implementation of the first aspect, based on engineering experience in passive thermal protection, the thickness of the expansion section profile and the thickness of the contraction section profile of the rocket mode Laval nozzle are taken into consideration.

[0023] According to a second aspect of the present invention, a dual-mode axisymmetric combined nozzle is provided, which is designed using the aforementioned dual-mode axisymmetric combined nozzle design method.

[0024] According to a third aspect of the present invention, a design apparatus for a dual-mode axisymmetric combined nozzle is provided, the combined nozzle comprising a rocket-mode Laval nozzle and an intake-mode Laval nozzle sleeved outside the rocket-mode Laval nozzle, the design apparatus comprising:

[0025] The acquisition module is used to acquire the initial expansion section profile and the wave-damping section profile of the reference nozzle. The initial expansion section profile is determined based on the analytical design method and the characteristic line method, and the wave-damping section profile is determined based on the maximum thrust theory. The reference nozzle refers to a single-channel Laval nozzle matched only when considering rocket modes.

[0026] The removal module is used to remove the approximately straight portion of the initial expansion section profile and use the remaining portion of the initial expansion section profile as the expansion section profile of the rocket mode Laval nozzle;

[0027] The first arc-drawing module is used to draw a straight line perpendicular to the reference nozzle axis from the end of the initial expansion section profile outwards, and draw a first arc upstream of the gas flow with a point on the straight line as the center. The first arc is used as the contraction section profile of the intake mode Laval nozzle. The first arc must satisfy the following: the first arc is connected to the outlet of the intake mode combustion chamber, and the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc is matched with the gas flow rate at the outlet of the intake mode combustion chamber when considering the thickness.

[0028] The second arc-drawing module is used to draw a second arc tangent to the wave-damping section profile downstream of the gas flow, with the point on the straight line as the center and the end of the first arc away from the outlet of the intake mode combustion chamber as the starting point. The portion of the wave-damping section profile before the tangent point is removed, and the remaining portion of the second arc and the wave-damping section profile is used as the expansion section profile of the intake mode Laval nozzle.

[0029] According to a fourth aspect of the present invention, an apparatus is provided, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the steps of the design method for a dual-modal axisymmetric combined nozzle.

[0030] According to a fifth aspect of the present invention, a computer-readable storage medium is provided, the computer-readable storage medium storing a computer program, which, when executed by a processor, implements the steps of the design method for a dual-modal axisymmetric combined nozzle.

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

[0032] This invention provides a design method for a dual-mode axisymmetric combined nozzle. By combining analytical design methods and maximum thrust theory in the design of the reference nozzle profile, the shape of the initial expansion section of the reference nozzle is controllable, and thrust is maximized while satisfying geometric constraints. Taking into account the size of the intake mode combustion chamber and the gas flow rate, the contraction section of the intake mode Laval nozzle is coupled into the nozzle design process. By removing part of the profile of the reference nozzle and using a smooth arc to connect the contraction and expansion sections of the outer bypass duct, the profile continuity of the expansion section of the inner duct Laval nozzle and the outer bypass duct Laval expansion section is ensured, reducing the interference of the outer bypass duct wall on the rocket gas and reducing the mass of the combined nozzle. The dual-mode axisymmetric combined nozzle of this invention can meet strict size constraints, has a high degree of dual-channel coupling, high profile continuity, and can effectively organize the flow field structure.

[0033] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, preferred embodiments are described below in detail with reference to the accompanying drawings. Attached Figure Description

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

[0035] Figure 1 This is a schematic diagram of the reference nozzle profile in the design method of a dual-modal axisymmetric combined nozzle of the present invention;

[0036] Figure 2 This is a schematic diagram of the combined nozzle design process in the design method of a dual-modal axisymmetric combined nozzle of the present invention;

[0037] Figure 3 This is a schematic diagram of the combined nozzle adjustment scheme in the design method of a dual-modal axisymmetric combined nozzle of the present invention.

[0038] In the diagram: 1 - Air-breathing mode combustion chamber; 2 - Rocket mode combustion chamber. Detailed Implementation

[0039] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0040] Combination Figure 1 and Figure 2 As shown, this invention provides a design method for a dual-mode axisymmetric combined nozzle, which includes a rocket-mode Laval nozzle and an intake-mode Laval nozzle fitted outside the rocket-mode Laval nozzle. The specific steps of the design method are as follows:

[0041] S1. Obtain the initial expansion section profile AB and the wave-damping section profile BF of the reference nozzle. The initial expansion section profile is determined based on analytical design methods and the method of characteristics, while the wave-damping section profile is determined based on maximum thrust theory. The reference nozzle refers to a single-channel Laval nozzle matched only when considering rocket modes.

[0042] Regarding the design of the initial expansion section profile of the reference nozzle, specifically, based on the method of characteristics, the initial expansion section profile of the reference nozzle is designed and obtained using fluid dynamics principles and computational software (such as CFD software). The initial expansion section profile of the reference nozzle must ensure that the combustion gases can smoothly accelerate and expand only in rocket mode, achieving the predetermined thrust efficiency.

[0043] In one embodiment, the process of determining the initial expansion section profile of the reference nozzle specifically includes:

[0044] The core point Mach number of the reference nozzle (i.e., point E) is determined using the method of characteristics. The coordinates of the end of the initial expansion section profile (i.e., the coordinates of point B) and the airflow angle β are iterated until the core point Mach number of the reference nozzle approaches the exit Mach number of the reference nozzle under the predetermined ideal state. Then, the initial expansion section profile of the reference nozzle is determined.

[0045] It should be noted that the coordinate system xOy corresponding to the coordinates at the end of the initial expansion section profile is a Cartesian coordinate system with the throat center of the reference nozzle as the origin, the x-axis perpendicular to the throat section and pointing to the nozzle exit, and the y-axis radially upward along the throat.

[0046] Specifically, when determining the core point Mach number of the reference nozzle, the basic principle of the method of characteristics is used. This involves solving for the core point Mach number using a unit process involving interior points, inverted wall points, and axis points. This Mach number is a key parameter in the design process, directly affecting the nozzle's expansion performance and thrust efficiency. To obtain the optimal initial expansion section profile, iterative calculations of the coordinates and airflow angle at the end of the initial expansion section profile are required. During the iteration process, the coordinates and airflow angle at the end of the initial expansion section profile are continuously adjusted until the core point Mach number of the reference nozzle approaches the predetermined ideal exit Mach number. It should be noted that the core point Mach number of the reference nozzle is optimal if it equals the predetermined ideal exit Mach number; if not, the closest value is considered optimal.

[0047] After the above iterative calculations, the initial expansion section profile of the reference nozzle can be obtained. This profile not only satisfies the basic principles of the method of characteristics, but also ensures that the nozzle can reach the predetermined exit Mach number under ideal conditions, thereby guaranteeing the thrust performance and efficiency of the nozzle.

[0048] In one embodiment, the method for determining the exit Mach number of the reference nozzle under ideal conditions is as follows:

[0049] a. Determine the inlet and outlet dimensions of the dual-mode axisymmetric combined nozzle based on the geometric constraints of the dual-mode axisymmetric combined nozzle.

[0050] In other words, firstly, based on design requirements and actual application scenarios, the inlet size (outlet diameter of intake mode combustion chamber 1) and outlet size (i.e., the final outlet diameter of the nozzle) of the dual-modal axisymmetric combined nozzle are determined. These dimensions are fundamental parameters in the design process and have a direct impact on the performance of the nozzle.

[0051] b. Determine the throat area of ​​the reference nozzle based on the exhaust gas flow rate of the rocket's mode combustion chamber.

[0052] In other words, based on the gas flow rate at the outlet of rocket mode combustion chamber 2, the throat area of ​​the reference nozzle (i.e., the rocket mode Laval nozzle) is calculated using the isentropic critical flow relation. This area needs to ensure that, in rocket mode, the gas can smoothly pass through the throat and accelerate its expansion.

[0053] For example, the formula for determining the throat area of ​​a reference nozzle is:

[0054]

[0055] in, γ is the gas flow rate at the outlet of the rocket's mode combustion chamber, γ is the specific heat ratio, and R is the gas constant. These represent the total temperature and total pressure at the reference nozzle inlet, A t-roc The throat area of ​​the reference nozzle.

[0056] c. Based on the throat area of ​​the reference nozzle and the exit size of the dual-mode axisymmetric combined nozzle, the exit Mach number of the reference nozzle under ideal conditions is determined using one-dimensional flow theory.

[0057] In other words, after determining the throat area of ​​the reference nozzle and the exit size of the dual-mode axisymmetric combined nozzle, the exit Mach number of the reference nozzle under ideal conditions can be calculated using one-dimensional flow theory. Specifically, the exit Mach number of the reference nozzle under ideal conditions can be obtained by solving the isentropic flow equation based on the throat area, exit area, and physical properties of the combustion gas (such as specific heat ratio, gas constant, etc.).

[0058] For example, the formula for determining the exit Mach number of a reference nozzle under ideal conditions is:

[0059]

[0060] Among them, A e Let A be the exit cross-sectional area of ​​the dual-mode axisymmetric combined nozzle. t-roc γ is the throat area of ​​the reference nozzle, γ is the specific heat ratio, and Ma0 is the exit Mach number of the reference nozzle under ideal conditions.

[0061] It should be understood that the axially symmetric combined nozzle has a circular cross-section along its axial direction, and the throat radius y of the reference nozzle can be calculated from the throat area of ​​the reference nozzle.* .

[0062] For example, regarding the determination of the initial expansion section profile of the quasi-nozzle, more specifically: combining the geometric constraints of the dual-mode axisymmetric combined nozzle and the exit Mach number of the reference nozzle under ideal conditions, calculate the coordinates of point B at the end of the initial expansion section profile and the airflow angle β. B Based on the coordinates of point B, the initial expansion section profile equation can be determined as follows:

[0063] y = y * +C1x 2 +C2x n

[0064] Among them, y * Given the throat radius of the reference nozzle, the equation coefficients C1 and C2 are determined by the following relationship:

[0065]

[0066] The x-coordinate of point B is determined by the following formula:

[0067]

[0068] The airflow angle at point B is determined by the following formula:

[0069]

[0070] Where ν1 is the Prandtl-Mayer angle corresponding to the exit Mach number of the reference nozzle under ideal conditions, and h is the radius of the exit of the reference nozzle.

[0071] Regarding the design of the shock wave profile of the reference nozzle, specifically, the profile is designed based on the maximum thrust theory. The main function of the shock wave profile is to reduce shock wave loss at the nozzle exit, thereby further improving thrust performance. The optimal shock wave profile is determined through numerical simulation and experimental verification.

[0072] In one embodiment, the process of determining the profile of the damping section of the reference nozzle specifically includes:

[0073] The position of the key point K of the reference nozzle is determined on the last right-hand characteristic line BE emanating from the end of the initial expansion section profile. Combining the maximum thrust theory, the position of the key point is iterated using the bisection method to make the exit size of the reference nozzle conform to the exit size constraint of the dual-mode axisymmetric combined nozzle. The wave-damping section profile of the reference nozzle is finally determined using the unit process of the change point method and the flow balance relationship.

[0074] More specifically, when determining the location of the key point of the reference nozzle, the last right-hand characteristic line is identified at the end of the initial expansion section profile of the reference nozzle. A key point needs to be determined on this last right-hand characteristic line (the location of this key point affects the shape of the wave-damping section profile and the performance of the nozzle). To accurately determine the location of the key point, a bisection method is used for iteration. During the iteration process, it is necessary to ensure that the exit size of the reference nozzle conforms to the exit size constraints of the dual-mode axisymmetric combined nozzle (achieved by adjusting the location of the key point). After determining the location of the key point, the change-point method element process is used to further determine the wave-damping section profile. In determining the wave-damping section profile, the flow balance relationship also needs to be considered. That is, the gas flow rate within the nozzle should remain constant at any cross-section. By comprehensively applying the change-point method element process and the flow balance relationship, the final wave-damping section profile of the reference nozzle is determined. The wave-damping section profile will ensure that the nozzle can generate sufficient thrust in rocket mode and meet the exit size constraints of the dual-mode axisymmetric combined nozzle.

[0075] For example, based on the maximum thrust theory, a point K is selected on the final rightward characteristic line. Using the nozzle length as a constraint, the bisection method is used to find point K that satisfies the length constraint at point F. The region BKF is solved using the change-point method element process and flow balance relationship to determine the profile of the choke section BF of the reference nozzle.

[0076] S2. Remove the approximately straight portion (i.e., GB segment) from the initial expansion segment profile, and use the remaining portion (i.e., AG segment) as the expansion segment profile of the rocket's modal Laval nozzle. This maintains the rocket exhaust flow in accordance with the source flow assumption while organically integrating the two channels.

[0077] Specifically, in the initial expansion section profile of the reference nozzle, approximately straight sections are identified and removed. These sections contribute little to nozzle performance improvement and increase overall weight. The remaining non-straight sections are used as the expansion section profile of the rocket-mode Laval nozzle, ensuring that the nozzle can efficiently convert combustion energy into thrust in rocket mode.

[0078] S3. Draw a straight line BO2 perpendicular to the axis of the reference nozzle from the end of the initial expansion section profile. With point O2 on the straight line as the center, draw a first arc JL upstream of the gas flow. Use the first arc JL as the contraction section profile of the intake mode Laval nozzle. The first arc JL must satisfy the following: the first arc JL is connected to the outlet of the intake mode combustion chamber 1, and the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc JL (i.e., the area of ​​the outer bypass annular section HM on the outlet section I-I of the rocket mode Laval nozzle) is matched with the gas flow rate at the outlet of the intake mode combustion chamber when considering the thickness.

[0079] In other words, a straight line perpendicular to the axis of the reference nozzle is drawn outward from the end of the initial expansion section profile of the reference nozzle. This straight line serves as the reference line for designing the contraction section profile of the intake mode Laval nozzle. Using a point on this line as the center (the location of this point needs to be determined based on the specific dimensions of the intake mode combustor outlet and the gas flow rate), a first arc is drawn upstream of the gas flow. The shape and position of the first arc must ensure that it connects to the outlet of the intake mode combustor, while simultaneously ensuring that the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc matches the gas flow rate at the intake mode combustor outlet.

[0080] In other words, air-breathing engines have a wide range of flow rate variations across a broad flight envelope. To ensure the flow capacity of the air-breathing mode Laval nozzle, the area of ​​the outer bypass duct annular section HM in the design state must match the operating point requirements of the exhaust gas flow rate at the air-breathing mode combustor outlet. Draw a straight line O2B along the positive y-axis from point B, and draw an arc LJ with point O2 as the center. This arc represents the contraction section of the air-breathing mode Laval nozzle, with a radius R. C The area of ​​the HM must be equal to the isentropic throat area corresponding to the exhaust gas flow rate at the intake mode combustion chamber outlet:

[0081]

[0082] in, γ is the exhaust gas flow rate at the outlet of the intake mode combustion chamber, γ is the specific heat ratio, and R is the gas constant. These represent the total temperature and total pressure at the inlet of the Laval nozzle in the intake mode, A t-tur Let HM be the area of ​​the annular cross-section of the outer duct.

[0083] S4. Taking point O3 on the straight line BO2 as the center, and taking the end of the first arc JL away from the outlet of the intake mode combustion chamber as the starting point, draw a second arc JN (with radius R) tangent to the profile of the wave-damping section downstream of the gas flow. D Remove the portion of the wave-damping section profile before the tangent point (i.e., the BN segment), and use the second arc JN and the remaining portion of the wave-damping section profile (i.e., the NF segment) as the expansion section profile of the intake mode Laval nozzle.

[0084] It should be understood that the second arc JN and the remaining part NF of the wave-damping section also serve as the expansion section of the rocket's mode Laval nozzle.

[0085] In one possible implementation, after obtaining the initial expansion section profile of the reference nozzle, the method further includes:

[0086] The contraction section profile of the reference nozzle is determined by combining the exit dimensions of the rocket mode combustion chamber. One end of the contraction section profile of the reference nozzle is connected to the exit of the rocket mode combustion chamber, and the other end smoothly transitions to the initial expansion section profile. The contraction section profile of the reference nozzle is used as the contraction section profile of the rocket mode Laval nozzle.

[0087] Specifically, analyzing the geometry (i.e., diameter) of the rocket's mode combustor 2 outlet, one end of the contraction section profile of the reference nozzle needs to connect with the rocket mode combustor outlet. Therefore, the starting point of the contraction section should be directly located on the boundary of the combustor outlet to ensure smooth gas flow into the nozzle. Starting from the starting point of the contraction section, a gradually contracting profile is designed until it smoothly transitions with the initial expansion section profile. At the connection between the contraction section profile and the initial expansion section profile, a smooth transition is required to ensure a smooth transition of the gas flow and reduce flow losses.

[0088] Since the reference nozzle is used as a single-channel Laval nozzle in rocket mode, its contraction section profile is also directly used as the contraction section profile of the rocket mode Laval nozzle.

[0089] For example, based on the applicable conditions of the Saul method, with ρ t ≥2r t Determining the radius of the arc of the contraction section of the rocket's modal Laval nozzle using the criterion ensures that the given solution is valid.

[0090] In one feasible approach, based on engineering experience with passive thermal protection, the thickness of the expansion section profile and the thickness of the contraction section profile of the rocket mode Laval nozzle are taken into account.

[0091] Specifically, in designing rocket mode Laval nozzles, the thickness of the nozzle profile not only affects its thermal protection performance but also directly relates to its mechanical strength and durability. Based on the operating environment and gas characteristics under rocket mode conditions, the thermal loads that the nozzle surface may bear are analyzed, including factors such as gas temperature, flow velocity, and the thermal conductivity of the nozzle material. Drawing on previous engineering experience in passive thermal protection design (empirical data typically comes from experimental testing, numerical simulations, or feedback from practical applications), the minimum required nozzle profile thickness under different thermal load conditions is determined.

[0092] It should be noted that when determining the thickness of the expansion and contraction sections of a rocket's modal Laval nozzle, factors such as the nozzle's thermal protection performance, mechanical strength, and durability must be comprehensively considered. If the initially determined thickness values ​​do not meet all requirements, adjustments and reassessment are necessary. Furthermore, factors such as the nozzle's manufacturing process and cost must also be considered. Excessively thick profiles may increase manufacturing costs; therefore, the thickness design should be optimized as much as possible while meeting performance requirements.

[0093] Combination Figure 3 As shown, this embodiment of the invention provides a dual-mode axisymmetric combined nozzle, designed using the aforementioned dual-mode axisymmetric combined nozzle design method. It should be understood that, because the rocket's mode combustion chamber operates in a high-temperature, high-pressure environment, and the rocket's mode Laval nozzle is located inside the engine, it is inconvenient to arrange adjustment mechanisms. Therefore, by moving the intake mode Laval nozzle back and forth (movement -Δx), the critical cross-sectional area of ​​the bypass channel can be easily adjusted, thereby controlling the gas flow rate in the intake mode and ensuring matching between the engine and exhaust system.

[0094] This invention provides a design apparatus for a dual-mode axisymmetric combined nozzle, the combined nozzle comprising a rocket-mode Laval nozzle and an air-breathing mode Laval nozzle sleeved outside the rocket-mode Laval nozzle. The design apparatus includes:

[0095] The acquisition module is used to acquire the initial expansion section profile and the wave-damping section profile of the reference nozzle. The initial expansion section profile is determined based on the characteristic line method, and the wave-damping section profile is determined based on the maximum thrust theory. The reference nozzle refers to a single-channel Laval nozzle matched only when considering rocket modes.

[0096] The removal module is used to remove the approximately straight portion of the initial expansion section profile and use the remaining portion of the initial expansion section profile as the expansion section profile of the rocket mode Laval nozzle.

[0097] The first arc-drawing module is used to draw a straight line perpendicular to the reference nozzle axis from the end of the initial expansion section profile. Taking a point on the straight line as the center, a first arc is drawn upstream of the gas flow. The first arc is used as the contraction section profile of the intake mode Laval nozzle. The first arc must satisfy the following: the first arc is connected to the outlet of the intake mode combustion chamber, and the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc, when considering the thickness, is matched with the gas flow rate at the outlet of the intake mode combustion chamber.

[0098] The second arc-drawing module is used to draw a second arc tangent to the wave-damping section profile downstream of the incoming gas flow, with the point on the straight line as the center and the end of the first arc away from the outlet of the combustion chamber in the intake mode as the starting point. The part of the wave-damping section profile before the tangent point is removed, and the remaining part of the second arc and the wave-damping section profile is used as the expansion section profile of the rocket mode Laval nozzle.

[0099] All relevant content regarding the steps involved in the aforementioned embodiment of a dual-modal axisymmetric combined nozzle design method can be referenced to the functional description of the corresponding functional module of the dual-modal axisymmetric combined nozzle design device in this embodiment of the invention, and will not be repeated here. The module division in this embodiment of the invention is illustrative and only represents a logical functional division. In actual implementation, there may be other division methods. Furthermore, the functional modules in the various embodiments of the invention can be integrated into a processor, exist as separate physical entities, or have two or more modules integrated into one module. The integrated modules described above can be implemented in hardware or as software functional modules.

[0100] In another embodiment of the present invention, a computer device is provided, comprising a processor and a memory. The memory stores a computer program, which includes program instructions. The processor executes the program instructions stored in the computer storage medium. The processor may be a Central Processing Unit (CPU), or other general-purpose processors, digital signal processors (DSPs), application-specific integrated circuits (ASICs), field-programmable gate arrays (FPGAs), or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, etc. It is the computing and control core of the terminal, suitable for implementing one or more instructions, specifically suitable for loading and executing one or more instructions in the computer storage medium to achieve a corresponding method flow or corresponding function. The processor described in this embodiment of the present invention can be used in the operation of a design method for a dual-modal axisymmetric combined nozzle.

[0101] In another embodiment of the present invention, a storage medium is provided, specifically a computer-readable storage medium (Memory), which is a memory device in a computer device used to store programs and data. It is understood that the computer-readable storage medium here can include both the built-in storage medium in the computer device and extended storage media supported by the computer device. The computer-readable storage medium provides storage space that stores the terminal's operating system. Furthermore, the storage space also stores one or more instructions suitable for loading and execution by a processor. These instructions can be one or more computer programs (including program code). It should be noted that the computer-readable storage medium here can be a high-speed RAM memory or a non-volatile memory, such as at least one disk storage device. The processor can load and execute one or more instructions stored in the computer-readable storage medium to implement the corresponding steps of the design method for a dual-modal axisymmetric combined nozzle in the above embodiments.

[0102] Those skilled in the art will understand that embodiments of the present invention can be provided as methods, systems, or computer program products. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention can take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0103] This invention is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0104] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0105] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0106] In this invention, the terms "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., refer to a specific feature, structure, material, or characteristic described in connection with that embodiment or example, which is included in at least one embodiment or example of the invention. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0107] Finally, it should be noted that the above-described embodiments are merely specific implementations of the present invention, used to illustrate the technical solutions of the present invention, and not to limit it. The scope of protection of the present invention is not limited thereto. Although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that any person skilled in the art can still modify or easily conceive of changes to the technical solutions described in the foregoing embodiments within the technical scope disclosed in the present invention, or make equivalent substitutions for some of the technical features; and these modifications, changes, or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention, and should all be covered within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. A design method for a dual-modal axisymmetric combined nozzle, characterized in that, The combined nozzle includes a rocket-mode Laval nozzle and an air-breathing mode Laval nozzle sleeved outside the rocket-mode Laval nozzle. The design method includes: The initial expansion section profile and the wave-damping section profile of the reference nozzle are obtained. The initial expansion section profile is determined based on the analytical design method and the characteristic line method, and the wave-damping section profile is determined based on the maximum thrust theory. The reference nozzle refers to a single-channel Laval nozzle matched only when considering rocket modes. Remove the approximately straight portion of the initial expansion section profile and use the remaining portion of the initial expansion section profile as the expansion section profile of the rocket mode Laval nozzle; Draw a straight line perpendicular to the reference nozzle axis from the end of the initial expansion section profile. With a point on the straight line as the center, draw a first arc upstream of the gas flow. Use the first arc as the contraction section profile of the intake mode Laval nozzle. The first arc must satisfy the following: the first arc connects to the outlet of the intake mode combustion chamber, and the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc, when considering the thickness, matches the gas flow rate at the outlet of the intake mode combustion chamber. Taking the point on the straight line as the center and the end of the first arc away from the outlet of the intake mode combustion chamber as the starting point, draw a second arc tangent to the wave-damping section profile downstream of the gas flow. Remove the part of the wave-damping section profile before the tangent point, and use the remaining part of the second arc and the wave-damping section profile as the expansion section profile of the intake mode Laval nozzle.

2. The design method for a dual-modal axisymmetric combined nozzle according to claim 1, characterized in that, The process of determining the initial expansion section profile of the reference nozzle specifically includes: The core point Mach number of the reference nozzle is determined by the method of characteristics. The coordinates and airflow angles at the end of the initial expansion section profile are iterated until the core point Mach number of the reference nozzle is close to the exit Mach number of the reference nozzle under the predetermined ideal state. Then the initial expansion section profile of the reference nozzle is determined.

3. The design method for a dual-modal axisymmetric combined nozzle according to claim 2, characterized in that, The method for determining the exit Mach number of the reference nozzle under ideal conditions includes: Based on the geometric constraints of the dual-mode axisymmetric combined nozzle, determine the inlet and outlet dimensions of the dual-mode axisymmetric combined nozzle; The throat area of ​​the reference nozzle is determined based on the exhaust gas flow rate at the rocket's modal combustion chamber outlet. Based on the throat area of ​​the reference nozzle and the exit size of the dual-mode axisymmetric combined nozzle, the exit Mach number of the reference nozzle under ideal conditions is determined using one-dimensional flow theory.

4. The design method of a dual-modal axisymmetric combined nozzle according to claim 3, characterized in that, The process of determining the profile of the bleed-suppression section of the reference nozzle specifically includes: The location of the key point of the reference nozzle is determined on the last rightward characteristic line emanating from the end of the initial expansion section profile. Combining the maximum thrust theory, the location of the key point is iterated using the bisection method to make the exit size of the reference nozzle conform to the exit size constraint of the dual-mode axisymmetric combined nozzle. The wave-damping section profile of the reference nozzle is finally determined using the unit process of the change point method and the flow balance relationship.

5. The design method of a dual-modal axisymmetric combined nozzle according to claim 1, characterized in that, After obtaining the initial expansion section profile of the reference nozzle, the following steps are also included: The contraction section profile of the reference nozzle is determined by combining the exit dimensions of the rocket mode combustion chamber. One end of the contraction section profile of the reference nozzle is connected to the exit of the rocket mode combustion chamber, and the other end smoothly transitions to the initial expansion section profile. The contraction section profile of the reference nozzle is used as the contraction section profile of the rocket mode Laval nozzle.

6. The design method of a dual-modal axisymmetric combined nozzle according to claim 5, characterized in that, Based on engineering experience in passive thermal protection, the thickness of the expansion section profile and the thickness of the contraction section profile of the rocket mode Laval nozzle are considered.

7. A dual-mode axisymmetric combined nozzle, characterized in that, The design method of the dual-modal axisymmetric combined nozzle as described in any one of claims 1 to 6 was used.

8. A design device for a dual-modal axisymmetric combined nozzle, characterized in that, The combined nozzle includes a rocket-mode Laval nozzle and an air-breathing mode Laval nozzle sleeved outside the rocket-mode Laval nozzle. The design device includes: The acquisition module is used to acquire the initial expansion section profile and the wave-damping section profile of the reference nozzle. The initial expansion section profile is determined based on the analytical design method and the characteristic line method, and the wave-damping section profile is determined based on the maximum thrust theory. The reference nozzle refers to a single-channel Laval nozzle matched only when considering rocket modes. The removal module is used to remove the approximately straight portion of the initial expansion section profile and use the remaining portion of the initial expansion section profile as the expansion section profile of the rocket mode Laval nozzle; The first arc-drawing module is used to draw a straight line perpendicular to the reference nozzle axis from the end of the initial expansion section profile outwards, and draw a first arc upstream of the gas flow with a point on the straight line as the center. The first arc is used as the contraction section profile of the intake mode Laval nozzle. The first arc must satisfy the following: the first arc is connected to the outlet of the intake mode combustion chamber, and the throat area formed by the expansion section profile of the rocket mode Laval nozzle and the first arc is matched with the gas flow rate at the outlet of the intake mode combustion chamber when considering the thickness. The second arc-drawing module is used to draw a second arc tangent to the wave-damping section profile downstream of the gas flow, with the point on the straight line as the center and the end of the first arc away from the outlet of the intake mode combustion chamber as the starting point. The portion of the wave-damping section profile before the tangent point is removed, and the remaining portion of the second arc and the wave-damping section profile is used as the expansion section profile of the intake mode Laval nozzle.

9. An apparatus comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements the steps of the design method for a dual-modal axisymmetric combined nozzle as described in any one of claims 1 to 6.

10. A computer-readable storage medium storing a computer program, characterized in that, When the computer program is executed by the processor, it implements the steps of the design method for a dual-modal axisymmetric combined nozzle as described in any one of claims 1 to 6.

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