A design method for multi-channel maximum thrust combined nozzles based on supersonic shear layer modeling and a supersonic shear layer modeling algorithm

Abstract

This invention first discloses a supersonic shear layer modeling algorithm, selecting airflow pressure and airflow angle as iterative variables. Based on the aerodynamic parameters of known points in the discrete flow field, it iterates continuously using a prediction-correction method until the relative error between iterative variables is less than a preset threshold, at which point the iterative variables converge. The supersonic shear layer points are solved based on this prediction-correction approach. Simultaneously, a multi-channel maximum thrust combined nozzle design method based on the supersonic shear layer modeling algorithm is also provided, including inverse design solutions for upstream channel I and channel II, flow field solution for the downstream mixing section of the initial surface, shear layer modeling solution, nozzle tail nozzle wall coordinate solution, maximum thrust nozzle control point solution, and maximum thrust nozzle control surface solution. This invention proposes a shear layer modeling algorithm based on the pressure-airflow angle balance concept, integrating the flow field structure into the nozzle design and eliminating complex flow structures in the flow field.

Description

Technical Field

[0001] This invention relates to the field of engine intake and exhaust system profile design technology, and mainly to a multi-channel maximum thrust combined nozzle design method and supersonic shear layer modeling algorithm based on supersonic shear layer modeling. Background Technology

[0002] Hypersonic flight generally refers to flight at a maximum cruise Mach number higher than 5. Sustained flight at such high speeds for extended periods places extremely high demands on the performance of the aircraft's propulsion system. Based on the variation of specific impulse performance of different engine types with flight Mach number, it is known that turbine engines possess high specific impulse performance at low Mach numbers. However, when the flight Mach number reaches 5 or higher, the specific impulse performance of turbine engines drops rapidly, making it difficult to meet flight requirements. Ramjet engines, on the other hand, offer better specific impulse performance. Furthermore, from an economic perspective, the development of hypersonic vehicle technology is increasingly trending towards reusability and horizontal takeoff and landing capabilities. Achieving a cruise speed of Mach 5 or higher from ground takeoff, while maintaining good specific impulse across such a wide speed range, is difficult to achieve with a single engine. Therefore, novel combined-cycle engines, which organically combine engines with good efficiency across different speed ranges, have become the optimal choice for hypersonic vehicle propulsion systems. Currently, combined-cycle engines under development mainly include rocket-based and turbine-based combined-cycle engines.

[0003] A rocket-based combined cycle (RBCC) engine is a novel type of combined cycle engine that combines an air-breathing high-speed engine with a rocket engine. Its general operating modes are as follows: it operates in rocket ejection mode at Mach 0-3, in subsonic ramjet mode at Mach 3-7, in scramjet mode at Mach 7-11, and in pure rocket mode when Mach > 11. A turbine-based combined cycle (TBCC) engine is a propulsion system that organically integrates a turbojet / turbofan engine or its improved versions with a ramjet engine. During low-speed flight, the turbojet / turbofan engine provides power, while at high Mach numbers, the ramjet engine provides power, thus providing continuous and efficient thrust to the aircraft across the entire thrust envelope. Compared to RBCCs, it does not require additional fuel, reducing its weight and increasing its range, while also offering the advantage of horizontal takeoff and landing.

[0004] Currently, many technical challenges have been encountered in the research of hypersonic combined cycle propulsion systems, and the design of the exhaust system considering complex flow conditions is one of the most critical aspects. Studies have shown that at Mach 6, the thrust provided by the exhaust system can account for more than 70% of the total thrust of the propulsion system. Other studies indicate that a 1% decrease in the thrust coefficient of the exhaust system leads to a 4% decrease in the net thrust of the engine, demonstrating the significant importance of exhaust system aerodynamic design for the performance of combined cycle engines. For parallel (vertical and internal / external parallel) combined cycle engine exhaust systems, although the turbine and ramjet engines have their own separate flow channels, the exhaust nozzle, as a shared component, needs to provide excellent thrust performance over a wide speed range. It must not only meet the performance requirements of both engines in their respective operating ranges but also fully consider the impact of imperfect geometric design and unreasonable aerodynamic parameter allocation on thrust performance, leading to supersonic airflow shearing between different flow channels, resulting in shock waves / boundary layer interference, flow separation, and other phenomena. Maximizing thrust at the high-speed cruise point is particularly crucial.

[0005] Therefore, this patent proposes a supersonic shear layer modeling algorithm and a multi-channel maximum thrust combined nozzle design method based on this algorithm. By rationally allocating the airflow parameters at the junction, the new algorithm is used to model the supersonic shear layer between the two channels, and on this basis, the maximum thrust nozzle profile is coupled for design, ensuring efficient flow and excellent thrust performance within the combined cycle engine exhaust system, providing sufficient thrust for hypersonic vehicles within a wide flight envelope. Summary of the Invention

[0006] Objective of the Invention: To address the problems existing in the background technology, this invention provides a multi-channel maximum thrust combined nozzle design method and a supersonic shear layer modeling algorithm based on supersonic shear layer modeling. By rationally allocating airflow parameters at the intersection, a new algorithm is used to model the supersonic shear layer between the two channels, and on this basis, the maximum thrust nozzle profile design is coupled to ensure efficient flow and excellent thrust performance within the combined cycle engine exhaust system, providing sufficient thrust for hypersonic vehicles within a wide flight envelope.

[0007] Technical solution: To achieve the above objectives, the technical solution adopted by this invention is as follows:

[0008] A supersonic shear layer modeling algorithm is proposed. An upstream supersonic airflow is formed at the initial value surface AB at the exit of channel I and the initial value surface B′O at the exit of channel II. A shear layer BMM′B′ is generated through the interaction of downstream channel III of channel I and downstream channel IV of channel II. The algorithm uses the known aerodynamic parameters of points 1, 2, 3, and 4 in the discrete flow field as input conditions and iteratively solves for the relevant aerodynamic parameters of downstream shear layer points 5 and 6. Specifically,

[0009] Points 1 and 3 are located in downstream channel III, and points 2 and 4 are located in downstream channel IV. Points 2 and 3 are shear layer points obtained using the algorithm in the previous step. AB, B′O, and the airflow intersection points B and B′ constitute the initial value surface of the algorithm. The modeled shear layer is BMM′B′, where BM and B′M′ are spatially coincident, but their airflow parameters are determined by channels III and IV, respectively. On BMM′B′, the airflow pressure and airflow angle are the same for two points with the same spatial location. Using airflow pressure and airflow angle as iterative variables, based on the aerodynamic parameters of known points in the discrete flow field, the prediction-correction method is used to iterate continuously until the relative error between the iterative variables obtained in the (n+1)th iteration and the nth iteration is less than a preset threshold. The solution is then complete, and the iterative variables converge.

[0010] Furthermore, the iterative process specifically includes:

[0011] First, based on the known aerodynamic parameters of points 1 and 3, the estimated values ​​of the spatial coordinates and airflow angles of points 5 and 6 are obtained, denoted as [x,y,θ]5 and [x,y,θ]6, respectively. Then, based on the aerodynamic parameters of points 2 and 4, the estimated values ​​of the airflow pressure of points 5 and 6 are obtained, denoted as p5 and p6, respectively. Based on the above estimated values, the aerodynamic parameters of points 1, 2, 3, and 4 in the next cycle are input for iteration, and the correction calculation is repeated until the aerodynamic parameters of points 5 and 6 meet the convergence condition.

[0012] A multi-channel maximum thrust combined nozzle design method employing the aforementioned supersonic shear layer modeling algorithm includes the inverse design solution of upstream channel I and channel II, with initial value surfaces AB and B′O as boundaries; the solution of downstream channels III and IV; the shear layer BMM′B′ modeling solution; the solution of tail nozzle wall surface AE coordinates; the solution of maximum thrust nozzle control point G; and the solution of maximum thrust nozzle control surface GE. Specifically,

[0013] Step L1: Perform surface design on upstream channel I and channel II using inverse design based on the method of rotational characteristic lines; using AB and B′O as initial calculation surfaces, obtain region I through inverse calculation using the interior point and axisymmetric point element process in the method of rotational characteristic lines. a II a The solution domain I is calculated along the flow direction of inlet HI and H2I2, with points K and K′. b II b When the parameters of the corresponding points K2 and K2′ are completely consistent with those of the boundary points K′ and K, and points K2′ and K2, then region I... b II b Solution complete; surfaces RK and R′K′, and surfaces R2K2 and R2′K2′ are two sets of surfaces that coincide in space and have identical parameters. By using region I... bII b After translating to the right until RK and R′K′ coincide, and R2K2 and R2′K2′ coincide, the flow field assembly parameters are matched and the flow field is assembled. The design of the surfaces HA and H2B′ of the upstream channel I and channel II is completed.

[0014] Step L2: Taking the initial value lines AB and B′O as boundaries, the overall flow field from upstream III... a Downward III b The downstream propulsion is calculated from the outside to the inside of downstream channels III and IV, starting from AB and B′O respectively. The flow field calculation uses the inward point and wall point element process in the method of characteristics, and the shear layer BMM′B′ is obtained by applying the supersonic shear layer modeling algorithm at the intersection. Taking the given nozzle geometric constraints and aerodynamic constraints as the target, i.e., the coordinates of point E or the inviscid aerodynamic parameters of point E, the control point G and the corresponding control surface equation GE are solved according to the maximum thrust theory. Within region III... b Solve for the directional curve AE based on the principle of flow conservation.

[0015] Beneficial effects:

[0016] (1) Compared with the existing design method of exhaust system of hypersonic combined cycle engine, the first step is to design a single-channel profile nozzle and open a second flow channel at a certain point on the nozzle wall. The relative position of the two channels is optimized through numerical simulation. This is a passive design method. Complex flows such as shear layer, shock wave, and expansion wave are very likely to appear in the flow field, which reduces the performance of the nozzle. This invention optimizes the calculation of airflow parameters at the intersection of the two channels and proposes a shear layer modeling algorithm based on the idea of ​​pressure-airflow angle balance. The flow field structure is integrated into the nozzle design, and the complex flow structure in the flow field is eliminated from the design concept.

[0017] (2) Based on the shear layer modeling algorithm, this invention successfully coupled the maximum thrust nozzle design method, further optimizing the nozzle performance on the basis of a simple flow field structure, and maximizing the thrust of the combined nozzle from the design perspective.

[0018] (3) This invention innovatively proposes a design system of upstream reverse design + shear layer modeling + downstream maximum thrust design, which provides a scientific and powerful design guide for the surface design of multi-channel combined nozzles of hypersonic combined cycle engines.

[0019] (4) The present invention can complete the axisymmetric / binary overall configuration and the design of dual / triple / multi-channel combined nozzles under the strict geometric aerodynamic constraints, according to the specific application object and needs. It has strong versatility and wide application. Attached Figure Description

[0020] Figure 1 This is a schematic diagram of the overall meridional structure of the dual-channel combined tail nozzle in an embodiment of the present invention;

[0021] Figure 2 This is a schematic diagram of the supersonic shearing layer modeling algorithm proposed in this invention;

[0022] Figure 3 This is an iterative flowchart of the supersonic shearing layer modeling algorithm proposed in this invention;

[0023] Figure 4 This is a theoretical distribution diagram of the airflow angle at the outlet of channel I in this embodiment of the invention;

[0024] Figure 5 This is a contour map of the Mach number of the inviscid ideal flow field in the embodiment of the present invention;

[0025] Figure 6 This is a contour map of the Mach number of the viscous flow field in the combined nozzle embodiment of the present invention. Detailed Implementation

[0026] The present invention will be further described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without inventive effort are within the scope of protection of the present invention.

[0027] This invention first proposes a supersonic shear layer modeling algorithm, based on Figure 1 The dual-channel combined nozzle has an overall meridional structure. Upstream supersonic airflow is formed at the initial value surface AB at the outlet of channel I and the initial value surface B′O at the outlet of channel II. A shear layer BMM′B′ is generated through the interaction of downstream channels III (channel I) and IV (channel II). The supersonic shear layer modeling algorithm uses the known aerodynamic parameters of points 1, 2, 3, and 4 in the discrete flow field as input conditions to iteratively solve for the relevant aerodynamic parameters of downstream shear layer points 5 and 6. Both upstream channels I and II have supersonic airflow at their outlets. At the airflow convergence point B and B′, in addition to equal pressure and equal airflow angle, other thermodynamic parameters are calculated based on the incoming flow parameters of their respective channels.

[0028] This algorithm is based on the supersonic rotational characteristic line method in gas dynamics. All aerodynamic parameters in the flow field are calculated using the compatibility equation, Bernoulli's equation, and the sound velocity equation. The compatibility equation, Bernoulli's equation, and the sound velocity equation are solved along the characteristic line difference grid as follows:

[0029]

[0030] ρVdV+dp=0

[0031] dp-a 2 dρ=0

[0032] Points 1 and 3 are located in downstream channel III, and points 2 and 4 are located in downstream channel IV. Points 2 and 3 are shear layer points already obtained using the algorithm in the previous step. Figure 2 As shown, AB, B′O and the airflow confluence points B and B′ constitute the initial value surface of the algorithm.

[0033] In this algorithm, since the airflow pressure and angle are the same at any point on the shear layer, airflow pressure and angle are selected as iterative variables. Based on the aerodynamic parameters of known points in the discrete flow field, an estimation-correction method is used for iterative calculation until the relative error between the (n+1)th and nth iterations is less than a preset threshold. At this point, the solution is complete, and the iterative variables converge. Specifically, as follows... Figure 3 As shown,

[0034] First, based on the known aerodynamic parameters of points 1 and 3, the slopes of lines 15 and 35 are solved, and then the spatial coordinates of point 5 (i.e., point 6) are obtained. The compatibility equation is solved along line 15, and the sound velocity equation and Bernoulli equation are solved along line 35, yielding the estimated values ​​of the airflow angle at point 5 (i.e., point 6), denoted as [x,y,θ]5 and [x,y,θ]6, respectively. Then, based on the aerodynamic parameters of points 2 and 4, the compatibility equation is solved along line 24, and the sound velocity equation and Bernoulli equation are solved along line 26, yielding the estimated values ​​of the airflow pressure at points 5 and 6, denoted as p5 and p6, respectively. Based on these estimated values, the aerodynamic parameters of points 1, 2, 3, and 4 in the next cycle are input for iteration, and the correction calculation is repeated until the aerodynamic parameters of points 5 and 6 meet the convergence condition.

[0035] Based on the aforementioned supersonic shear layer modeling algorithm, this invention also provides a multi-channel maximum thrust combined nozzle design method, such as... Figure 1 As shown, the entire design method includes the inverse design solution of upstream channels I and II, with initial value surfaces AB and B′O as boundaries; the solution of downstream flow field channels III and IV, with initial value surfaces AB and B′O as boundaries; the modeling solution of the shear layer BMB′M′; the solution of the nozzle tail nozzle wall coordinates; the solution of the maximum thrust nozzle control point G; and the solution of the maximum thrust nozzle control surface GE. Specifically,

[0036] Step L1: Perform surface design on upstream channel I and channel II using inverse design based on the method of rotational characteristic lines; using AB and B′O as initial calculation surfaces, obtain region I through inverse calculation using the interior point and axisymmetric point element process in the method of rotational characteristic lines. a II a The solution domain I is calculated along the flow direction of inlet HI and H2I2, with points K and K′. b II b When the parameters of the corresponding points K2 and K2′ are completely consistent with those of the boundary points K′ and K, and points K2′ and K2, then region I...b II b Solution complete; surfaces RK and R′K′, and surfaces R2K2 and R2′K2′ are two sets of surfaces that coincide in space and have identical parameters. By using region I... b II b After translating to the right until RK and R′K′ coincide, and R2K2 and R2′K2′ coincide, the flow field assembly parameters are matched and the flow field is assembled. The design of the surfaces HA and H2B′ of the upstream channel I and channel II is completed.

[0037] Step L2: Taking the initial value lines AB and B′O as boundaries, the overall flow field from upstream III... a Downward III b The downstream propulsion is calculated from the outside to the inside of downstream channels III and IV, starting from AB and B′O respectively. The flow field calculation uses the inward point and wall point element process in the method of characteristics, and the shear layer BMM′B′ is obtained by applying the supersonic shear layer modeling algorithm at the intersection. Taking the given nozzle geometric constraints and aerodynamic constraints as the target, i.e., the coordinates of point E or the inviscid aerodynamic parameters of point E, the control point G and the corresponding control surface equation GE are solved according to the maximum thrust theory. Within region III... b Solve for the directional curve AE based on the principle of flow conservation.

[0038] The supersonic shear layer modeling method and the maximum thrust combined nozzle design method based on this algorithm proposed in this invention can design dual / triple / multi-channel combined nozzles for axisymmetric / binary configurations, given the inlet conditions of each channel, the distribution of airflow confluence surface parameters and the geometric / aerodynamic constraints of the nozzle.

[0039] This invention essentially provides a technical framework for a hypersonic combined cycle engine multi-channel nozzle, focusing on its profile design. The core design principles are a given converging airflow profile, upstream nozzle reverse design, shear layer modeling algorithm, and downstream nozzle maximum thrust design, with the shear layer modeling algorithm as the core technology. This design framework is applicable to axisymmetric / binary configurations for dual / triple / multi-channel combined nozzle design and can be embedded in all flow channel profile designs involving supersonic shear layers, exhibiting wide applicability and excellent robustness. Specific embodiments are given below to further demonstrate the technical effects of this invention.

[0040] With a linear airflow distribution of 12 degrees at the outlet of channel I as the design condition, and a design Mach number of 2.24 at its outlet, and a horizontally uniform airflow distribution at the outlet of channel II, and a design Mach number of 1.8 at its outlet, the profile design of the dual-channel maximum thrust combined nozzle is carried out using this method.

[0041] like Figure 4 As shown, the theoretical distribution of the outlet airflow angle of channel I almost completely coincides with the actual result of the outlet airflow angle designed in this invention. The Mach numbers of the outlet airflow of channel I and channel II are as follows: Figure 5 As shown, the results agree well with the design values, and the overall results indicate that the shear layer modeling algorithm and combined nozzle design method proposed in this invention are effective; the viscous flow field of the combined nozzle is as follows: Figure 6 As shown, its axial thrust coefficient reaches C fx =0.9624, the combined tail nozzle has excellent performance, which proves that the design method proposed in this invention can enable the tail nozzle to provide the engine with excellent thrust performance.

[0042] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A supersonic shear layer modeling method, characterized by, An upstream supersonic flow is formed by the initial value surface AB at the exit of the channel I and the initial value surface B'O at the exit of the channel II, and a shear layer BMM'B' is generated under the interaction of the downstream channel III of the channel I and the downstream channel IV of the channel II; the supersonic shear layer modeling method uses the known aerodynamic parameters of points 1, 2, 3 and 4 in the discrete flow field as input conditions, and iteratively solves the related aerodynamic parameters of points 5 and 6 in the downstream shear layer; specifically, Points 1 and 3 are distributed in the downstream channel III, points 2 and 4 are distributed in the downstream channel IV, and points 2 and 3 are the shear layer points obtained by using the method in the previous step; AB, B'O and airflow intersection points B and B' constitute the initial value surface of the method; the modeled shear layer is BMM'B', wherein BM and B'M' have the same spatial position but the airflow parameters are determined by the channel III and the channel IV respectively; on BMM'B', the airflow pressure and the airflow angle of two points with the same spatial position are the same, the airflow pressure and the airflow angle are used as the iterative variables, based on the aerodynamic parameters of the known points in the discrete flow field, the method of prediction-correction is used for continuous iteration until the relative error between the iterative variables solved in the n+1th cycle and the n th cycle is less than a preset threshold, the solving is completed, and the iterative variables converge; The iterative process specifically includes: First, the estimated values of the spatial coordinates and the airflow angle of points 5 and 6 are solved according to the aerodynamic parameters of points 1 and 3, and are represented as [x, y, q]5 and [x, y, q]6 respectively; then the estimated values of the airflow pressure of points 5 and 6 are solved according to the aerodynamic parameters of points 2 and 4, and are represented as p5 and p6 respectively; based on the above estimated values, the aerodynamic parameters of points 1, 2, 3 and 4 are input into the next cycle for iteration, and the correction calculation is repeated until the aerodynamic parameters of points 5 and 6 meet the convergence condition.

2. A method for designing a multi-passage maximum thrust combined nozzle using the supersonic shear layer modeling method of claim 1, characterized in that, The reverse design solving of the upstream channel I and the channel II with the initial value surfaces AB and B'O as the boundary, the solving of the downstream channel III and the channel IV, the modeling solving of the shear layer BMM'B', the solving of the wall surface AE coordinates of the tail nozzle, the solving of the control point G of the maximum thrust nozzle and the solving of the control surface GE of the maximum thrust nozzle; specifically, Step L1, using inverse design based on the method of characteristic curve with rotation to upstream channel I, channel II profile design; with AB, B'O as the initial value surface, through the process of internal point and axisymmetric point element in the method of characteristic curve with rotation to get region I a , II a and the corresponding point K, K', along the flow direction to calculate the region I b , II b and the corresponding point K2 and K2', until the boundary point K' and K, point K2' and K2 parameter is completely consistent, region I b , II b solving is completed; surface RK and R'K', surface R2K2 and R2'K2' are two groups of surfaces that are coincident in space, and are completely the same in parameters, by shifting region I b , II b to the right to RK and R'K', surface R2K2 and R2'K2' coincide after completing the flow field assembly parameter matching, and then the flow field assembly is carried out, completing the profile HA, H2B' design of upstream channel I, channel II; Step L2, with initial value surface AB, B'O as boundary, the flow field as a whole from upstream Ⅲ a downstream Ⅲ b propulsion, respectively from AB, B'O start to downstream passage Ⅲ, passage Ⅳ from outside to inside calculation, flow field calculation using the inner point, wall point element process in the method of characteristics, and at the intersection, the supersonic shear layer modeling method is used to solve and obtain the shear layer BMM'B'; with the given nozzle geometric constraint and aerodynamic constraint as the target, according to the maximum thrust theory, the control point G and the corresponding control surface equation GE are solved, and in the region Ⅲ b According to the flow conservation equation, the tail nozzle wall AE is solved.

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