CN103678774B - Designing method for supersonic velocity thrust exhaust nozzle considering inlet parameter unevenness - Google Patents

Abstract

The invention discloses a supersonic thrust nozzle design method considering the non-uniform inlet parameters, which includes the following steps: (1) using the swirling characteristic line method to determine the initial value line according to the inlet parameter distribution of the thrust nozzle to be designed distribution area; according to the inlet parameters of the thrust nozzle to be designed, the design pressure ratio and the selected asymmetry factor G, the initial expansion angle of the upper wall surface and the initial expansion angle of the lower wall surface of the thrust nozzle to be designed at the sharp point of the throat are respectively determined. expansion angle; (2) first determine the coordinates and flow field parameters of each discrete point at the sharp point of the throat of the thrust nozzle to be designed, so as to obtain the flow parameters at the sharp point of the throat, and then determine the core area of the thrust nozzle to be designed The flow parameters of all the characteristic lines; (3) Using the wave elimination method to determine the upper and lower wall curves of the thrust nozzle to be designed, the design of the thrust nozzle to be designed can be completed. Therefore, the present invention can design a thrust nozzle considering the non-uniformity of inlet parameters, and produce better thrust performance.

Description

Design method of supersonic thrust nozzle considering non-uniform inlet parameters

technical field

The invention relates to a design method of a supersonic thrust nozzle considering non-uniform inlet parameters, and belongs to the technical field of supersonic exhaust nozzles.

Background technique

As the core technology of air-breathing hypersonic vehicles, scramjet technology has gradually become a research hotspot in various countries. The scramjet engine is composed of an inlet, an isolation section, a combustion chamber and an exhaust nozzle. As an important part of the scramjet engine, the main function of the tail nozzle is to fully expand the high-enthalpy airflow generated by the combustion chamber to generate as high a thrust as possible, while taking into account the requirements of aircraft aerodynamic balance such as lift and pitching moment, the performance of the tail nozzle It has a great influence on the performance of the aircraft and has always been an important field of hypersonic research.

In the actual work of the scramjet engine, whether it is the compression shock wave system of the intake port to the high-speed free flow, the oscillation of the asymmetric shock wave train in the isolation section, or the mixing of air and fuel in the combustion chamber, oscillating combustion, etc., Both will cause non-uniform distribution of flow field parameters at nozzle inlet. In addition, the tail nozzle of the scramjet has no constriction section and geometric throat, which cannot effectively rectify the airflow like the usual Laval nozzle, and is directly connected to the combustion chamber, so the inlet airflow of the tail nozzle is inevitable has considerable inhomogeneity.

At home and abroad, limited research has been carried out on the influence of non-uniform inlet airflow on the aerodynamic performance of scramjet tail nozzle. Snelling conducted a numerical simulation on the non-uniform inlet of hypersonic vehicle tail nozzle, and believed that the non-uniform inlet would increase the thrust of the aircraft and reduce the overall moment. Schindel used two jets with different Mach numbers to simulate the non-uniform inlet of the nozzle, and expanded isentropically with the uniform inlet airflow to the same ambient pressure respectively, compared the outlet momentum of the two, and concluded that the non-uniform inlet airflow velocity distribution caused the thrust performance of the nozzle It is generally concluded that the decline will not exceed 1%. Goel used numerical simulation to study the influence of different non-uniform distribution of inlet parameters on the performance of the nozzle. The results showed that the performance of the nozzle had a great relationship with the non-uniform distribution of the inlet. Kushida used the one-dimensional mixed flow processing method, and estimated that the impact of non-uniform inlet on the nozzle is about 1%. Ebrahimi believes that the non-uniformity of the inlet has a certain influence on the pressure distribution in the front section of the nozzle, and believes that the non-uniformity has no more than 2% influence on the thrust of the nozzle.

Le Jialing, Wang Xiaodong et al. used numerical simulation methods to study the influence of inlet temperature profile on the flow field structure of scramjet tail nozzle, and the results showed that temperature non-uniformity had little influence on the nozzle. Xu Jinglei, Quan Zhibin et al. conducted experiments and numerical studies on the influence of Mach number non-uniform inlet on tail nozzle performance, and the results showed that non-uniform inlet caused tail nozzle thrust performance to decrease, negative lift to increase and pitching moment to decrease .

The above series of research results show that the non-uniformity of the tail nozzle inlet airflow will have a certain impact on the tail nozzle flow field structure and aerodynamic performance based on uniform parameter design. Then, can the non-uniformity of the inlet air flow be considered at the beginning of the design of the nozzle, so as to obtain a nozzle profile that matches the inlet non-uniform air flow? In this regard, there is no research report so far, only Richard at the NASA Langley Research Center in the United States designed a supersonic wind tunnel nozzle considering the non-uniformity of the inlet airflow by using the method of swirling characteristic lines. There is no relevant research on the design of the thrust nozzle considering the non-uniform inlet, but it is urgently needed for the performance research of the scramjet tail nozzle under the real inlet condition.

Contents of the invention

The present invention aims at the defect that there is no special design method for the thrust nozzle of the detached ramjet engine considering the non-uniform inlet at present, and proposes a design method for the thrust nozzle considering the non-uniform inlet parameters. The method considers the inlet airflow Mach number along the height direction Under the premise of non-uniform distribution, the isentropic curve of scramjet asymmetric nozzle is designed by using the swirling characteristic line. The performance difference between nozzles designed with and without consideration of inlet non-uniformity was further studied under the same non-uniform inlet conditions; therefore, the present invention can design thrust nozzles considering inlet parameter non-uniformity, and produce relatively Good thrust performance.

For realizing above technical purpose, the present invention will take following technical scheme:

A method for designing a supersonic thrust nozzle considering non-uniform inlet parameters, comprising the following steps: (1) using the method of swirling characteristic lines to determine the distribution area of the initial value line according to the distribution of the inlet parameters of the thrust nozzle to be designed; in addition , according to the inlet parameters of the thrust nozzle to be designed, the design pressure ratio and the selected asymmetry factor G, the initial expansion angle of the upper wall surface and the initial expansion angle of the lower wall surface of the thrust nozzle to be designed at the sharp point of the throat are respectively determined ; The asymmetric factor G is the ratio of the initial expansion angle of the upper and lower walls at the sharp point of the throat of the thrust nozzle to be designed; (2) According to the thrust nozzle to be designed determined in step (1) at the throat The initial expansion angle of the upper wall surface and the initial expansion angle of the lower wall surface at the sharp point of the upper wall, first determine the coordinates and flow field parameters of each discrete point at the sharp point of the throat of the thrust nozzle to be designed, in order to obtain the thrust nozzle to be designed at the throat The flow parameters at the sharp point, and then use the method of rotating characteristic lines, according to the distribution area of the initial value line obtained in step (1), determine the flow parameters of all the characteristic lines of the core area of the thrust nozzle to be designed, the flow The parameters include pressure, temperature, velocity, and airflow direction angle; (3) Using the wave elimination method, according to the flow parameters of each characteristic line in the core area of the thrust nozzle to be designed in step (2), determine the thrust nozzle to be designed The design of the thrust nozzle to be designed can be completed through the curves of the upper and lower walls.

The iterative formula of the spin characteristic line method is:

in: is the abscissa, is the ordinate, the origin of the coordinates is at the lower corner of the nozzle inlet, the x direction is the horizontal direction of the inlet, and the y direction is the normal direction of the horizontal direction. is the local flow direction angle, is the local flow Mach number and >1, for the local mobile Mach angle, is the flow type parameter, for two-dimensional flow =0, for axisymmetric flow =1; P is the pressure at a specific point of the nozzle to be designed; V is the inlet velocity of the nozzle to be designed;

, , They are the coordinate values of three different points on the characteristic line unit. Initially, they are the coordinate values of the three different points on the initial value line. Coordinate values of different points; is the coordinate value of the corresponding discrete point on the feature line to be sought; is the ordinate , average of; is the ordinate , average of.

The sharp point of the throat of the thrust nozzle to be designed is arranged in a transitional arc.

The radius of the transition arc is 10% of the nozzle inlet height.

There is an angle α between the inlet airflow and the horizontal direction; and the length of the upper wall of the thrust nozzle to be designed is greater than the length of the lower wall.

The initial expansion angle of the upper wall surface of the thrust nozzle to be designed at the sharp point of the throat , the initial expansion angle of the lower wall are determined by the following formula:

,

in: is the Prandtl-Meier expansion angle corresponding to the gas flow expanding from the nozzle inlet to the design outlet Mach number; is the angle between the inlet airflow of the nozzle and the horizontal axis; G is the ratio of the initial expansion angle of the upper and lower walls at the sharp point of the throat of the thrust nozzle to be designed.

According to above technical scheme, with respect to prior art, the present invention has following advantage:

The invention can design the thrust nozzle considering the non-uniformity of the inlet parameters, and can produce better thrust performance.

Description of drawings

Figure 1 shows the geometric parameters of the asymmetric thrust nozzle.

Figure 2 shows the expansion wave of an asymmetric thrust nozzle at the sharp point of the throat.

Figure 3 shows the nozzle wave system of an asymmetric thrust nozzle.

Figure 4 is a schematic diagram of wave dissipation on the wall of an asymmetric nozzle.

Fig. 5 is the Mach number distribution curve at the nozzle inlet of Embodiment 1.

Fig. 6 is the Mach number distribution curve at the nozzle inlet of Embodiment 2.

Fig. 7 is the total temperature distribution curve of the nozzle inlet of embodiment 2.

detailed description

The accompanying drawings disclose, without limitation, the structural schematic diagrams of the preferred embodiments involved in the present invention; the technical solution of the present invention will be described in detail below in conjunction with the accompanying drawings.

As shown in Figure 1, according to the asymmetric thrust nozzle design method considering the non-uniform inlet parameters of the present invention, given the inlet parameters of the nozzle and the design pressure ratio, it is first necessary to determine the sharp points of the throat on the upper and lower walls The initial expansion angle at . As shown in Fig. 1, the acute angle of the throat of the upper and lower walls is determined by the geometric control parameters G and and the exit Mach number is determined, is the angle between the inlet airflow of the nozzle and the horizontal axis, G is the ratio of the initial expansion angle of the upper and lower walls, that is, the acute angle of the throat, , is the initial expansion angle of the upper wall, is the initial expansion angle of the lower wall.

,

is the Prandtl-Meier expansion angle corresponding to the gas flow expanding from the nozzle inlet to the design outlet Mach number. Inlet airflow direction angle limited to 0 to between, When and When the nozzle appearance is similar. For all asymmetric shapes, it is assumed in this paper that the length of the upper wall is greater than the length of the lower wall. For the scramjet engine, the airflow out of the combustion chamber is not completely along the direction of the nozzle axis, but has a certain angle. The previous design method cannot take into account the angle of the inlet airflow, but the design in this paper The method can consider the airflow direction at the inlet of the nozzle, so it is very advantageous.

Second, determine the parameters of the flow field at the sharp point of the throat. After determining the initial expansion angle, it is necessary to determine the flow parameters at the throat tip. Considering the adverse effect of the sharp point of the throat on the flow field, and the fact that the sharp point of the throat cannot be a theoretical sharp point in actual processing, but must be a circular arc, so the sharp point of the throat is replaced by a small arc. The design shows that when the arc radius is 10% of the inlet height, the length of the nozzle increases by less than 1%, so it is feasible to replace the sharp point of the throat with a small arc. If the throat is a cusp, it is a singularity from which an expansion fan emanates (see Figure 2). After being replaced by arcs, the arcs corresponding to the initial expansion angles of the throats on the upper and lower walls of the nozzle are discretized. In this way, after discretization, the coordinates and airflow direction angles of each point on the arc of the throat are determined, so that the flow parameters of each point on the arc can be determined by the Prandtl-Meyer relation.

After the coordinates and flow parameters of each point in the throat are determined, the airflow direction angle and airflow turning angle on all characteristic lines in the core area can be determined. There are mainly interior point element processes and direct wall point element processes to determine the flow in the core region. For the case where G is not equal to 1, assuming that the length of the lower wall is smaller than that of the upper wall, the expansion wave emitted at the entrance cusp of the upper wall will be reflected on the lower wall. Since the design of the nozzle requires that there are no shock waves and compression waves in the nozzle, the corresponding section of the lower wall must be a straight line. The reflected wave on the lower wall intersects with the expansion wave emitted from behind the cusp on the upper wall, until a certain reflected wave interacts with the last expansion wave on the upper cusp and the Mach number is equal to the design Mach number (point g in Figure 3), the entire core area Flow OK.

When the present invention determines the flow field parameters, it is determined by the swirl characteristic line method. For the swirl characteristic line method (references: Maurice J. Zucrow and Joe D. Hoffman, Gas Dynamics [M]. John Wiley & Sons. Inc ), the selected initial value line area is determined according to the inlet parameter distribution of the thrust nozzle to be designed; at the same time, each discrete point on the initial value line is determined by the flow parameters and coordinate points at the inlet of the thrust nozzle to be designed Express.

The iterative formula of the spin characteristic line method is:

in: is the abscissa, is the ordinate, the origin of the coordinates is at the lower corner of the nozzle inlet, the x direction is the horizontal direction of the inlet, and the y direction is the normal direction of the horizontal direction. is the local flow direction angle, is the local flow Mach number and >1, for the local mobile Mach angle, is the flow type parameter, for two-dimensional flow =0, for axisymmetric flow =1;

, , They are the coordinate values of three different points on the characteristic line unit. Initially, they are the coordinate values of the three different points on the initial value line. Coordinate values of different points; is the coordinate value of the corresponding discrete point on the feature line to be sought; is the ordinate , average of; is the ordinate , average of.

Finally, determine the nozzle wall points. After determining the flow in the core region, in two-dimensional flow, the wall ac can be determined according to the wave dissipation at the wall (reference: B. M. Argrow and G. Emanuel, Comparison of Minimum Length Nozzles. Journal of Fluids Engineering). Figure 4 shows the determination process of the wall point. The positions of points a and b and related aerodynamic parameters should have been known through the previous steps. c is the wall point to be obtained. Firstly, the geometric sum of point c is obtained through the streamline equation of point a and the left-hand characteristic line equation of point b. Aerodynamic parameters, and then modify the parameters of c by matching the flow rate on the characteristic line of segment ab and the characteristic line of segment bc. By repeating the above steps, the accurate parameter of point c can be obtained.

Example 1:

The area ratio of the nozzle is 4.69, the ratio of the length of the upper expansion surface to the height of the throat is 10.58, and the ratio of the length of the lower wall to the height of the throat is 3.7. Figure 5 shows the distribution curve of the inlet Mach number. In the case of a nozzle drop pressure ratio of 60, the thrust of the nozzle obtained by this design method is 2.8% higher than that obtained by the conventional method.

Example 2:

The area ratio of the nozzle is 2.58, the ratio of the length of the upper expansion surface to the height of the throat is 7.7, and the ratio of the length of the lower wall to the height of the throat is 4. Attached drawing 6 is the inlet Mach number distribution curve, and attached drawing 7 is the inlet total temperature distribution. curve. In the case of a nozzle drop pressure ratio of 50, the thrust of the nozzle obtained by this design method is 1% higher than that obtained by the conventional method.

Claims (6)

1. a kind of consideration intake condition supersonic speed heterogeneous propelling nozzle method for designing is it is characterised in that the method is considering Inlet Mach number, along on the premise of short transverse non-uniform Distribution, is started using there being the rotation method of characteristic curves to devise ultra-combustion ramjet The constant entropy molded line of machine offset nozzle, specifically includes following steps: (1), using there being the rotation method of characteristic curves, sprays according to thrust to be designed The intake condition distribution of pipe, determines the distributed areas of initial value line；In addition, according to the intake condition of propelling nozzle to be designed, setting Meter pressure ratio and selected asymmetric factor g, determine this propelling nozzle to be designed at the beginning of the upper wall surface at throat's cusp respectively Beginning divergence cone angle, lower wall surface initial bubble angle；Asymmetric factor g is the upper and lower wall at throat's cusp of this propelling nozzle to be designed The ratio at the initial bubble angle in face；(2) upper wall at throat's cusp according to the propelling nozzle to be designed determined in step (1) Face initial bubble angle, lower wall surface initial bubble angle, first determine each discrete point coordinates at throat's cusp of propelling nozzle to be designed And flow field parameter, to obtain flow parameter at throat's cusp for the propelling nozzle to be designed, and then using having the rotation method of characteristic curves, root The distributed areas of the initial value line obtaining according to step (1), by interior dot element process or direct wall dot element process, determine The flow parameter of all characteristic curves of propelling nozzle core space to be designed, described flow parameter includes airflow direction angle and air-flow Deflection angle；(3) adopt wave absorption method, the flowing ginseng of each characteristic curve according to the described propelling nozzle core space to be designed of step (2) Number, determines the upper and lower wall curve of propelling nozzle to be designed, you can complete the design of propelling nozzle to be designed.

2. according to claim 1 consider intake condition supersonic speed heterogeneous propelling nozzle method for designing it is characterised in that Described have rotation the method for characteristic curves iterative formula be:

Wherein: x is abscissa, y is vertical coordinate, and in the lower angle point of nozzle inlet, x direction is import horizontal direction to zero, y Direction is the normal direction of horizontal direction；θ is local flow direction angle, and for local flowing Mach number and m > 1, μ is the local Mach that flows to m Angle, δ is pattern of flow parameter, for two-dimensional flow δ=0, for axial symmetry flow δ=1；P is jet pipe to be designed in certain bits Pressure at point；V is the inlet velocity of jet pipe to be designed；

(x₁,y₁)、(x₂,y₂)、(x₃,y₃) be respectively three differences on characteristic curve unit coordinate figure, be initial value when initial The coordinate figure of three differences on line, later then according to have on the rotation characteristic curve of back asked for of the method for characteristic curves three The coordinate figure of difference；(x₄,y₄) it is corresponding discrete point coordinate figure on characteristic curve to be asked；It is vertical coordinate y₁、y₄Meansigma methodss；It is vertical coordinate y₂、y₄Meansigma methodss.

3. according to claim 1 consider intake condition supersonic speed heterogeneous propelling nozzle method for designing it is characterised in that It is in that transition arc is arranged at throat's cusp of described propelling nozzle to be designed.

4. according to claim 3 consider intake condition supersonic speed heterogeneous propelling nozzle method for designing it is characterised in that The radius of described transition arc is the 10% of nozzle inlet height.

5. according to claim 1 consider intake condition supersonic speed heterogeneous propelling nozzle method for designing it is characterised in that There is angle α between described inlet air flow and horizontal direction；And the upper wall surface length of propelling nozzle to be designed is more than lower wall surface Length.

6. according to claim 1 consider intake condition supersonic speed heterogeneous propelling nozzle method for designing it is characterised in that Upper wall surface initial bubble angle δ at throat's cusp for the described propelling nozzle to be designed_u, lower wall surface initial bubble angle δ_lRespectively by under Formula determines:

δ_u=ν_e/2·(1-α/ν_e), δ_l=α+g ν_e/2·(1-α/ν_e)

Wherein: ν_eIt is corresponding Prandtl-Meyer divergence cone angle when nozzle inlet expand into design exit Mach number of air-flow；α is Angle between nozzle inlet air-flow and trunnion axis；G is the first of the upper and lower wall at throat's cusp of this propelling nozzle to be designed The ratio of beginning divergence cone angle.