CN114329822A - Design method of multi-channel maximum thrust combined spray pipe based on supersonic shear layer modeling and supersonic shear layer modeling algorithm - Google Patents
Design method of multi-channel maximum thrust combined spray pipe based on supersonic shear layer modeling and supersonic shear layer modeling algorithm Download PDFInfo
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Abstract
The invention firstly discloses a modeling algorithm of an ultrasonic shear layer, which selects airflow pressure and airflow angle as iteration variables, continuously iterates by adopting a pre-estimation-correction method based on the pneumatic parameters of known points in a discrete flow field until the relative error between the iteration variables is smaller than a preset threshold value, and the iteration variables are converged; solving the supersonic shear layer point based on the estimation-correction idea; the design method comprises inverse design solving of an upstream channel I and a channel II, initial value surface downstream mixed section flow field solving, shear layer modeling solving, nozzle tail nozzle wall surface coordinate solving, maximum thrust nozzle control point solving and maximum thrust nozzle control surface solving; the invention provides a shear layer modeling algorithm based on the idea of pressure-airflow angle balance, and a flow field structure is integrated into a spray pipe design, so that a complex flow structure in the flow field is eliminated.
Description
Technical Field
The invention relates to the technical field of profile design of an engine intake and exhaust system, in particular to a design method of a multi-channel maximum thrust combined spray pipe based on supersonic shear layer modeling and a supersonic shear layer modeling algorithm.
Background
Hypersonic flight generally refers to flight procedures with a maximum cruise mach number above 5, where prolonged sustained flight at such high speeds places extremely high demands on the performance of the dynamic propulsion system of the aircraft. According to the change situation of the specific impulse performance of different types of engines along with the flight Mach number, the turbine engine can have higher specific impulse performance when flying at a low Mach number, when the flight Mach number reaches more than 5, the specific impulse performance of the turbine engine is rapidly reduced, the flight requirement is difficult to meet, and the ramjet engine has better specific impulse performance. Meanwhile, due to the economic consideration, the technical development of the hypersonic flight vehicle gradually tends to be reusable and capable of horizontal take-off and landing. In order to realize cruising from ground takeoff to the maximum speed of more than Mach 5, the cruising with better specific impulse performance in the wide speed range is difficult to realize only by a single engine, so that a novel combined cycle engine which organically combines the engines with better efficiency in different speed ranges becomes a poor choice for a propulsion system of a hypersonic aircraft. The combined cycle engines currently under development mainly comprise rocket-based and turbine-based combined cycle engines.
A Rocket-Based Combined Cycle (RBCC) is a new Combined Cycle engine combining an air-breathing high-speed engine with a Rocket engine. The general working mode is as follows: when Ma is 0-3, the rocket motor works in a rocket ejection mode, when Ma is 3-7, the rocket motor works in a sub-combustion stamping mode, when Ma is 7-11, the rocket motor works in a super-combustion stamping mode, and when Ma is more than 11, the rocket motor works in a pure rocket mode; the turbo Based Combined Cycle engine (TBCC) is a propulsion system organically Combined with a pressure engine on the basis of the existing mature turbojet/turbofan engine or its modified version, and is powered by the turbojet/turbofan engine during low-speed flight and powered by a ramjet engine during high-mach-number flight, so that the TBCC can provide continuous and efficient thrust for an aircraft in a full envelope range. Compared with RBCC, it does not need to carry extra fuel, has lightened self weight, and the voyage is farther, still possesses the advantage of level take-off and landing simultaneously.
Many technical challenges are currently encountered in the development of hypersonic combined propulsion systems, of which the design of exhaust systems considering complex flows is the most critical part. It has been found that at a flight mach number of 6, the thrust provided by the exhaust system can be more than 70% of the total thrust of the propulsion system. Additional studies have indicated that every 1% reduction in the thrust coefficient of an exhaust system results in a 4% reduction in the engine installed net thrust, and it can be seen that the aerodynamic design of the exhaust system is of great importance to the performance of a combined cycle engine. For a parallel-connection (up-down parallel connection and inside-outside parallel connection) combined cycle engine exhaust system, although a turbine engine and a ramjet engine have independent runners, a tail nozzle is used as a shared component of the turbine engine and the ramjet engine, and needs to provide excellent thrust performance in a wide speed domain range, so that not only are the performance requirements of the two engines in respective working intervals met, but also the influences of phenomena such as mutual shearing of supersonic airflows among airflows of different runners, further induced shock wave/boundary layer interference, flow separation and the like on the thrust performance caused by incomplete geometric design and unreasonable pneumatic parameter distribution are fully considered, and particularly the maximization of the thrust can be realized at a high-speed cruise point.
Therefore, the patent discloses an ultrasonic shear layer modeling algorithm and a design method of a multi-channel maximum thrust combined nozzle based on the algorithm. By reasonably distributing airflow parameters at the intersection, modeling an ultrasonic shear layer between two channels by using a new algorithm and coupling the design of the maximum thrust nozzle profile on the basis, efficient flow and excellent thrust performance in an exhaust system of the combined cycle engine are ensured, and sufficient thrust is provided for the hypersonic aerocraft in a wide flight envelope.
Disclosure of Invention
The purpose of the invention is as follows: aiming at the problems in the prior art, the invention provides a design method of a multi-channel maximum thrust combined spray pipe based on supersonic shear layer modeling and a supersonic shear layer modeling algorithm.
The technical scheme is as follows: in order to achieve the purpose, the invention adopts the technical scheme that:
a supersonic shear layer modeling algorithm is characterized in that an upstream supersonic air flow is formed on an initial value surface AB of an outlet of a channel I and an initial value surface B ' O of an outlet of a channel II, and a shear layer BMM ' B ' is generated under the interaction of a downstream channel III of the channel I and a downstream channel IV of the channel II; the modeling algorithm of the supersonic shear layer adopts the known pneumatic parameters of 1, 2, 3 and 4 points in the discrete flow field as input conditions, and iteratively solves the related pneumatic parameters of the downstream shear layer points 5 and 6; in particular, the amount of the solvent to be used,
the points 1 and 3 are distributed in a downstream channel III, the points 2 and 4 are distributed in a downstream channel IV, and the two points 2 and 3 are shear layer points which are obtained in the last step length by using the algorithm; AB. B 'O and the airflow junction B, B' form an initial value surface of the algorithm; the modeled shear layer is BMM 'B', wherein BM and B 'M' are spatially coincident, but the gas flow parameters are determined by channel III and 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 iteration variables, and iteration is continuously carried out by adopting a prediction-correction method based on the pneumatic parameters of the known points in the discrete flow field until the relative error between the iteration variables solved by the n +1 th cycle and the n-th cycle is less than a preset threshold value, so that the solution is completed, and the iteration variables are converged.
Further, the iterative process specifically includes:
firstly, the space coordinates and gas of 5 and 6 points are obtained by solving according to the known pneumatic parameters of 1 and 3 pointsThe predicted values of the flow angles, denoted x, y, theta, respectively]5And [ x, y, theta ]]6(ii) a Then, solving and obtaining the estimated values of the air pressure at the two points 5 and 6 according to the pneumatic parameters at the two points 2 and 4, which are respectively expressed as p5 and p 6; and inputting the pneumatic parameters of 1, 2, 3 and 4 points in the next cycle for iteration based on the estimated values, and repeating the correction calculation until the pneumatic parameters of 5 and 6 points meet the convergence condition.
A design method of a multi-channel maximum thrust combined nozzle adopting the supersonic shear layer modeling algorithm comprises the steps of solving the inverse design of an upstream channel I and a channel II which are divided by initial value surfaces AB and B ' O, solving the inverse design of a downstream channel III and a channel IV, solving the modeling of a shear layer BMM ' B ', solving a tail nozzle wall surface AE coordinate, solving a maximum thrust nozzle control point G and solving a maximum thrust nozzle control surface GE; in particular, the amount of the solvent to be used,
l1, performing profile design on the upstream channel I and the upstream channel II by using inverse design based on a spiral characteristic line method; respectively obtaining the region I by taking AB and B' O as initial calculation value surfaces and performing reverse calculation through the unit process of the inner point and the axisymmetric point in the rotation characteristic line methoda，ⅡaAnd point K, K2At the inlet HI, H2I2Along-flow calculation solving area Ib，ⅡbCorresponding point K2And K2', up to boundary points K' and K, point K2' and K2When the parameters of (A) are completely consistent, region Ib，ⅡbCompleting solution; faces RK and R 'K', face R2K2And R2′K2' two groups of planes with identical parameters spatially coincident with each other are formed by dividing region Ib，ⅡbTranslate rightward to RK and R 'K', plane R2K2And R2′K2After superposition, flow field assembly parameters are matched to carry out flow field assembly, and molded surfaces HA and H of an upstream channel I and a channel II are finished2B' designing;
l2, taking the initial lines AB and B' O as boundaries, the whole flow field is from upstream IIIaDown IIIbThe downstream channel III and the downstream channel IV are respectively calculated from the outside to the inside by the aid of the travel propulsion from AB to B' O, and the flow field is calculated to enable the downstream channel III and the downstream channel IV to be in the insideSolving and obtaining a shear layer BMM 'B' by using the inner point and wall surface point unit processes in the characteristic line method and applying the ultrasonic shear layer modeling algorithm at the intersection; using the given geometrical constraint and pneumatic constraint of the nozzle as targets, i.e. point E coordinate or point E non-viscous pneumatic parameter, solving control point G and corresponding control surface equation GE according to maximum thrust theory, and III in the regionbAnd solving the turning profile AE according to the flow conservation.
Has the advantages that:
(1) compared with the design method of the existing hypersonic combined cycle engine exhaust system, the design method is characterized in that a single-channel molded surface spray pipe is designed, a second flow channel is arranged at a certain position on the wall surface of the spray pipe, the relative positions of the two channels are optimized through numerical simulation, the passive design method is adopted, complex flows such as shear layers, shock waves, expansion waves and the like are easy to occur in the flow field, and the performance of the spray pipe is reduced.
(2) The invention successfully couples a maximum thrust jet pipe design method based on a shear layer modeling algorithm, further optimizes the jet pipe performance on the basis of simple flow field structure, and maximizes the thrust of the combined jet pipe from the design point of view.
(3) The invention innovatively provides a design system of upstream reverse design, shear layer modeling and downstream maximum thrust design, and provides a scientific and powerful design guiding idea for the profile design of the multi-channel combined nozzle of the hypersonic combined cycle engine.
(4) The invention can complete the design of the axisymmetric/binary overall configuration and the double/triple/multi-channel combined spray pipe under the condition of meeting the rigorous geometric pneumatic constraint according to specific application objects and requirements, and has strong universality and wide application.
Drawings
FIG. 1 is a schematic view of an integral meridian plane structure of a two-channel combined exhaust nozzle in an embodiment of the invention;
FIG. 2 is a schematic diagram of a modeling algorithm for a supersonic shear layer proposed by the present invention;
FIG. 3 is an iterative flow chart of a modeling algorithm for a supersonic shear layer proposed by the present invention;
FIG. 4 is a theoretical distribution of the exit flow angle of channel I in an embodiment of the present invention;
FIG. 5 is a contour plot of the Mach number of the inviscid ideal flow field of the composite nozzle of an embodiment of the present invention;
FIG. 6 is a contour plot of the viscous flow field Mach number for a composite nozzle in an embodiment of the present invention.
Detailed Description
The present invention will be further described with reference to the accompanying drawings. It is to be understood that the embodiments described are only a few embodiments of the present invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The invention firstly provides an ultrasonic shear layer modeling algorithm, which is based on a double-channel combined tail nozzle integral meridian plane structure shown in figure 1, wherein an upstream ultrasonic air flow is formed on an initial value surface AB of an outlet of a channel I and an initial value surface B ' O of an outlet of a channel II, and a shear layer BMM ' B ' is generated under the interaction of a downstream channel III of the channel I and a downstream channel IV of the channel II; the supersonic shear layer modeling algorithm adopts the known pneumatic parameters of 1, 2, 3 and 4 points in the discrete flow field as input conditions, and iteratively solves the related pneumatic parameters of the downstream shear layer points 5 and 6. The outlets of two runners of the upstream channel I and the channel II are both supersonic air flows, the point B and the point B' at the intersection of the air flows meet the condition that the pressure is equal and the air flow angle is equal, and other thermodynamic parameters are calculated according to the incoming flow parameters of the runners.
The algorithm is calculated based on a supersonic velocity rotating characteristic line method in gas dynamics, and all pneumatic parameters in a flow field are calculated through a compatibility equation, a Bernoulli equation and a sound velocity equation. Respectively solving a compatibility equation, a Bernoulli equation and a sound velocity equation along the characteristic line differential grid as follows:
ρVdV+dp＝0
dp-a2dρ＝0
the points 1 and 3 are distributed in the downstream channel III, the points 2 and 4 are distributed in the downstream channel IV, and the two points 2 and 3 are shear layer points which are obtained by solving the shear layer points by using the algorithm in the last step length, as shown in FIG. 2; AB. B 'O and the flow junction B, B' form the initial surface of the algorithm.
In the algorithm, because the airflow pressure and the airflow angle at any point on the shear layer are the same, the airflow pressure and the airflow angle are selected as iteration variables, the iteration is continuously carried out by adopting a prediction-correction method based on the pneumatic parameters of the known point in the discrete flow field until the relative error between the iteration variables solved by the n +1 th cycle and the n-th cycle is less than a preset threshold value, the solution is completed, and the iteration variables are converged. As shown in particular in figure 3 of the drawings,
firstly, solving the slopes of a line 15 and a line 35 according to the known aerodynamic parameters of 1 point and 3 points, further solving to obtain a 5-point (namely 6-point) space coordinate, solving a compatibility equation along the line 15, solving a sound velocity equation and a Bernoulli equation along the line 35 to obtain a pre-estimated value of the airflow angle of the 5-point (namely 6-point), which is respectively expressed as [ x, y, theta ] theta]5And [ x, y, theta ]]6(ii) a Then, according to the pneumatic parameters of the two points 2 and 4, solving a compatibility equation along a line 24, solving a sound velocity equation and a Bernoulli equation along a line 26, and solving to obtain predicted values of the air pressure of the two points 5 and 6, which are respectively expressed as p5 and p 6; and inputting the pneumatic parameters of 1, 2, 3 and 4 points in the next cycle for iteration based on the estimated values, and repeating the correction calculation until the pneumatic parameters of 5 and 6 points meet the convergence condition.
Based on the supersonic shear layer modeling algorithm, the invention also provides a design method of the multi-channel maximum thrust combined nozzle, as shown in figure 1, the whole design method comprises an upstream channel I taking initial value surfaces AB and B 'O as boundaries, inverse design solution of a channel II, an AB and B' O downstream flow field channel III, a channel IV solution, shear layer BMB 'M' modeling solution, nozzle tail nozzle wall surface AE coordinate solution, maximum thrust nozzle control point G solution and maximum thrust nozzle control surface GE solution. In particular, the amount of the solvent to be used,
l1, performing profile design on the upstream channel I and the upstream channel II by using inverse design based on a spiral characteristic line method; respectively obtaining the region I by taking AB and B' O as initial calculation value surfaces and performing reverse calculation through the unit process of the inner point and the axisymmetric point in the rotation characteristic line methoda，ⅡaAnd point K, K2At the inlet HI, H2I2Along-flow calculation solving area Ib，ⅡbCorresponding point K2And K2', up to boundary points K' and K, point K2' and K2When the parameters of (A) are completely consistent, region Ib，ⅡbCompleting solution; faces RK and R 'K', face R2K2And R2′K2' two groups of planes with identical parameters spatially coincident with each other are formed by dividing region Ib，ⅡbTranslate rightward to RK and R 'K', plane R2K2And R2′K2After superposition, flow field assembly parameters are matched to carry out flow field assembly, and molded surfaces HA and H of an upstream channel I and a channel II are finished2B' designing;
l2, taking the initial lines AB and B' O as boundaries, the whole flow field is from upstream IIIaDown IIIbPerforming travel propulsion, namely calculating a downstream channel III and a downstream channel IV from outside to inside from AB and B ' O respectively, calculating a flow field by using an inner point and wall surface point unit process in a characteristic line method, and solving and obtaining a shear layer BMM ' B ' by applying the ultrasonic shear layer modeling algorithm at a junction; using the given geometrical constraint and pneumatic constraint of the nozzle as targets, i.e. point E coordinate or point E non-viscous pneumatic parameter, solving control point G and corresponding control surface equation GE according to maximum thrust theory, and III in the regionbAnd solving the turning profile AE according to the flow conservation.
According to the supersonic shear layer modeling method and the maximum thrust combined spray pipe design method based on the algorithm, double/triple/multichannel combined spray pipe design can be performed according to axisymmetric/binary configuration under the conditions of inlets of all channels, parameter distribution of an airflow intersection surface and geometric/pneumatic constraint of the spray pipe.
The invention essentially provides a technical framework which takes a hypersonic combined cycle engine multichannel tail nozzle as an object, takes the profile design as a target, takes a given cross airflow profile, an upstream nozzle inverse design, a shear layer modeling algorithm and a downstream nozzle maximum thrust design as a core design idea, and takes the shear layer modeling algorithm as a core technology. The design framework is suitable for axisymmetric/binary configuration, carries out double/triple/multichannel combined spray pipe design, can be embedded into all runner profile designs comprising an ultrasonic shear layer, and has wide applicability and excellent robustness. Specific examples are given below to further demonstrate the technical effects of the present invention.
Taking the linear distribution of the airflow angle at the outlet of the channel I as 12 degrees as a design condition, and setting the design Mach number of the outlet of the channel I to be 2.24; and the airflow at the outlet of the channel II is horizontally and uniformly distributed, the Mach number of the outlet of the channel II is given to be 1.8, and the profile design of the double-channel maximum thrust combined spray pipe is carried out by the method.
As shown in fig. 4, the theoretical distribution of the exit flow angles of channel i almost completely coincides with the actual results of the exit flow angles contemplated by the present invention. The Mach numbers of the outlet air flows of the channel I and the channel II are shown in FIG. 5 and are well matched with the design value, and the comprehensive result shows that the shear layer modeling algorithm and the combined nozzle design method provided by the invention are effective; the viscous flow field of the combined nozzle is shown in figure 6, and the axial thrust coefficient reaches CfxThe combined exhaust nozzle has excellent performance equal to 0.9624, and proves that the design method provided by the invention can enable the exhaust nozzle to provide excellent thrust performance for an engine.
The above description is only of the preferred embodiments of the present invention, and it should be noted that: it will be apparent to those skilled in the art that various modifications and adaptations can be made without departing from the principles of the invention and these are intended to be within the scope of the invention.
Claims (3)
1. A modeling algorithm of an ultrasonic shear layer is characterized in that an upstream ultrasonic air flow is formed on an initial value surface AB of an outlet of a channel I and an initial value surface B ' O of an outlet of a channel II, and a shear layer BMM ' B ' is generated under the interaction of a downstream channel III of the channel I and a downstream channel IV of the channel II; the modeling algorithm of the supersonic shear layer adopts the known pneumatic parameters of 1, 2, 3 and 4 points in the discrete flow field as input conditions, and iteratively solves the related pneumatic parameters of the downstream shear layer points 5 and 6; in particular, the amount of the solvent to be used,
the points 1 and 3 are distributed in a downstream channel III, the points 2 and 4 are distributed in a downstream channel IV, and the two points 2 and 3 are shear layer points which are obtained in the last step length by using the algorithm; AB. B 'O and the airflow junction B, B' form an initial value surface of the algorithm; the modeling shear layer is BMM 'B', wherein BM and B 'M' are coincident in spatial position, but the gas flow parameters are respectively determined by a channel III and a channel IV; 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 iteration variables, and iteration is continuously carried out by adopting a prediction-correction method based on the pneumatic parameters of the known points in the discrete flow field until the relative error between the iteration variables solved by the n +1 th cycle and the n-th cycle is less than a preset threshold value, so that the solution is completed, and the iteration variables are converged.
2. The supersonic shear layer modeling algorithm of claim 1, wherein the iterative process specifically comprises:
firstly, according to the known pneumatic parameters of two points 1 and 3, the estimated values of space coordinates and air flow angle of two points 5 and 6 are obtained by solving, and are respectively expressed as [ x, y, theta ]]5And [ x, y, theta ]]6(ii) a Then, solving and obtaining the estimated values of the air pressure at the two points 5 and 6 according to the pneumatic parameters at the two points 2 and 4, which are respectively expressed as p5 and p 6; and inputting the pneumatic parameters of 1, 2, 3 and 4 points in the next cycle for iteration based on the estimated values, and repeating the correction calculation until the pneumatic parameters of 5 and 6 points meet the convergence condition.
3. The design method of the multi-channel maximum thrust combined nozzle adopting the supersonic shear layer modeling algorithm of any one of claims 1-2 is characterized by comprising the steps of solving the inverse design of an upstream channel I and a channel II which are demarcated by initial value surfaces AB and B ' O, solving a downstream channel III and a channel IV, solving the modeling of a shear layer BMM ' B ', solving a tail nozzle wall surface AE coordinate, solving a maximum thrust nozzle control point G and solving a maximum thrust nozzle control surface GE; in particular, the amount of the solvent to be used,
l1, performing profile design on the upstream channel I and the upstream channel II by using inverse design based on a spiral characteristic line method; respectively obtaining the region I by taking AB and B' O as initial calculation value surfaces and performing reverse calculation through the unit process of the inner point and the axisymmetric point in the rotation characteristic line methoda，ⅡaAnd point K, K2At the inlet HI, H2I2Along-flow calculation solving area Ib，ⅡbCorresponding point K2And K2', up to boundary points K' and K, point K2' and K2When the parameters of (A) are completely consistent, region Ib，ⅡbCompleting solution; faces RK and R 'K', face R2K2And R2′K2' two groups of planes with identical parameters spatially coincident with each other are formed by dividing region Ib，ⅡbTranslate rightward to RK and R 'K', plane R2K2And R2′K2After superposition, flow field assembly parameters are matched to carry out flow field assembly, and molded surfaces HA and H of an upstream channel I and a channel II are finished2B' designing;
l2, taking the initial lines AB and B' O as boundaries, the whole flow field is from upstream IIIaDown IIIbPerforming travel propulsion, namely calculating a downstream channel III and a downstream channel IV from outside to inside from AB and B ' O respectively, calculating a flow field by using an inner point and wall surface point unit process in a characteristic line method, and solving and obtaining a shear layer BMM ' B ' by applying the ultrasonic shear layer modeling algorithm at a junction; using the given geometrical constraint and pneumatic constraint of the nozzle as targets, i.e. coordinate of point E or E-point inviscid pneumatic parameters, solving control point G and corresponding control surface equation GE according to maximum thrust theory, and in region IIIbAnd solving the turning profile AE according to the flow conservation.
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