CN108304601B - Method for judging transition of boundary layer of hypersonic aircraft - Google Patents
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Abstract
The invention provides a method for judging transition of a boundary layer of a hypersonic aircraft, which is used for determining the transition position of the surface of the aircraft based on early-stage test data; acquiring a laminar flow field corresponding to the orthogonalization of the aircraft surface in a test state, and acquiring N value distribution of the aircraft surface by adopting an improved e-N method considering a transverse flow mode; determining the N value N of triggering transition by combining the aircraft surface transition position determined by the test data0And is used for judging the transition position of the target aircraft. The method is based on test data, the three-dimensional characteristics of a flow field of the booster body/multiplier body external high supersonic speed aircraft are considered, the disturbance waves of a second mode and a lower frequency transverse flow mode are considered when the e-N method is adopted for integration, and the transition criterion N for accurately predicting the transition position is obtained0The method provides an effective judgment method for the accurate prediction of the boundary layer transition of the lifting body/wave rider external form hypersonic aircraft.
Description
Technical Field
The invention belongs to the technical field of aircraft boundary layer stability analysis and transition prediction, particularly relates to a method for judging transition of a boundary layer of a hypersonic aircraft, and particularly relates to a method for judging transition of a boundary layer of a booster/multiplier hypersonic aircraft based on flight/test data.
Background
During flight test of a hypersonic aircraft (hereinafter referred to as an aircraft for short), two flow states of laminar flow and turbulent flow exist on the surface of the aircraft. The differences in flow regimes result in differences in aircraft lift, drag, surface heat flow, engine performance, and inlet launch performance. The flow process from laminar flow to turbulent flow is defined as transition in the present invention. The accurate prediction of the transition position of the surface of the aircraft is the key of the design of the aircraft and determines the success or failure of the flight test.
The method for predicting the transition position of the surface of the aircraft comprises the following steps: the method comprises a traditional e-N method based on semi-experience of a linear stability theory, a method (PSE) for solving a parabolic stability equation, Direct Numerical Simulation (DNS), a turbulence transition mode and an engineering transition criterion, wherein the e-N method based on the stability theory is most widely applied in engineering application. In the field of the aviation industry, the e-N method is generally considered to be the most effective method for predicting the transition position.
However, with the conventional e-N method, small perturbations of various frequencies exist within the boundary layer, which start to grow in amplitude at different locations as they propagate downstream; the method can calculate the accumulated increase multiple of disturbance waves with various frequencies in the downstream direction from the position where the disturbance waves start to increase by using the linear stability theory, wherein the increase multiple reaches a certain value N firstly0The perturbation wave causes a transition.
The concrete expression is as follows:
or
The e-N method considers the relative increase of the disturbance amplitude, does not consider the actual magnitude of the vibration amplitude, and triggers the transition N0The magnitude of the value needs to be given by experiment or experience.
For the transition of the boundary layer of the hypersonic aircraft, the second mode disturbance wave is generally considered to play a dominant role, and the e-N method only integrates the disturbance wave in the frequency band where the second mode is located during integration. After reviewing the transition problem of hypersonic speed, Saric gives an overviewN of high-speed two-dimensional flow trigger transition0The value is about 10, which is currently substantially the norm. However, for hypersonic aircrafts with complex shapes such as a lifting body/a waverider and the like, due to the fact that the three-dimensional property of a flow field is strong and the transverse flow is obvious, under the condition, the early-given trigger transition N010 is no longer applicable. Therefore, if the transition position of the surface of the hypersonic aircraft is to be accurately obtained, how to obtain the accurate N triggering transition0The value is a problem to be solved.
Disclosure of Invention
The invention aims to overcome the defects of the prior art and provides a transition position judgment method of a hypersonic aircraft boundary layer, which introduces a cross flow influence into the judgment of the transition position based on an improved e-N method and gives how to obtain N triggering transition in the e-N method considering the cross flow influence0The problem of engineering application of the traditional e-N method is solved.
The technical solution of the invention is as follows:
a method for judging transition of a boundary layer of a hypersonic aircraft comprises the following steps:
step 1, determining a transition position of the surface of the early-stage aircraft;
1.1, acquiring early-stage test data of the aircraft, wherein a required test state is selected according to requirements;
1.2 determining the transition position of the surface of the aircraft in the test state according to the test data;
step 2, obtaining transition criterion N by adopting an improved e-N stability analysis and judgment method0；
2.1 calculating the laminar flow field in the test state, and interpolating the flow field to a grid which is normal and orthogonal to the surface of the aircraft to obtain an orthogonal flow field;
2.2 analyzing a neutral curve by taking the orthogonalized flow field as a basic flow based on the orthogonalized flow field to obtain a neutral curve considering the influence of transverse flow;
2.3 selecting a second mode disturbance wave and a cross flow mode disturbance wave with amplitude growth rate larger than 0 based on the obtained neutral curve, integrating the disturbance waves along the direction of potential flow by adopting an e-N method, and taking the envelope curve of the calculated N values of all the disturbance waves as the N value distribution of the surface of the aircraft;
2.4 based on the transition position of the aircraft surface and the N-value distribution of the aircraft surface, determine the N-value transition criterion N for triggering transition0；
Step 3, enabling the obtained transition criterion N0And judging the surface transition position of the target aircraft.
According to the judgment method for the transition e-N stability analysis of the hypersonic velocity boundary layer, the transition position of the surface of the aircraft is determined according to the early-stage test data; acquiring a laminar flow field corresponding to the orthogonalization of the aircraft surface in a test state, and acquiring N value distribution of the aircraft surface by adopting an improved e-N method considering a transverse flow mode; determining the N value N of triggering transition by combining the aircraft surface transition position determined by the test data0. The method is based on test data, the three-dimensional characteristic of a flow field of the booster body/multiplier body external high supersonic speed aircraft is considered, because the flow field is possibly in a cross flow mode, the invention adopts the e-N method for integrating, besides considering the disturbance wave in the second mode, also considering the disturbance wave in the cross flow mode with lower frequency, further obtaining a transition judgment analysis method considering the cross flow mode, and obtaining the transition criterion N for accurately predicting the transition position0The method provides an effective judgment method for the accurate prediction of the boundary layer transition of the lifting body/wave rider external form hypersonic aircraft.
Drawings
The accompanying drawings, which are included to provide a further understanding of the embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 is a flowchart illustrating a method for determining transition of a boundary layer of a hypersonic aircraft according to an embodiment of the present invention;
FIG. 2 is a schematic diagram illustrating a transition position determination of the early aircraft according to the present invention;
FIG. 3 is a comparison graph of neutral curves of the hypersonic aircraft transition position prediction method with and without consideration of cross flow;
FIG. 4 is a diagram showing N-value distribution of disturbance waves of different frequencies along the direction of potential flow in the method for judging transition of a boundary layer of a hypersonic aircraft provided by the invention;
FIG. 5 is a schematic diagram illustrating distribution of N values in the method for predicting a transition position of a hypersonic aircraft according to the present invention;
FIG. 6 shows the method for predicting the transition position of hypersonic flight vehicle of the present invention0Determining a schematic diagram;
in fig. 6, a black line indicates a transition position.
Detailed Description
Specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings. In the following description, for purposes of explanation and not limitation, specific details are set forth in order to provide a thorough understanding of the present invention. However, it will be apparent to one skilled in the art that the present invention may be practiced in other embodiments that depart from these specific details.
It should be noted that, in order to avoid obscuring the present invention with unnecessary details, only the device structures and/or processing steps that are closely related to the scheme according to the present invention are shown in the drawings, and other details that are not so relevant to the present invention are omitted.
Referring to fig. 1, the embodiment provides a method for judging transition of a boundary layer of a hypersonic aircraft, which is implemented by the following steps:
the method comprises the following steps: a transition position determination method based on flight test data;
step 1.1 hypersonic aircraft test data;
in the step, the experimental data is based on the previous aircraft surface heat flow measurement experimental data, the data can be obtained by adopting the technology known in the field, the experimental state corresponding to the experimental data acquisition is selected according to the requirement, the experimental data can be adjusted according to the actual requirement,
step 1.2, determining the surface transition position of the aircraft;
determining the transition position of the aircraft surface in the test state by the heat flow test data in the step 1.1 and combining a heat flow value (Q) obtained by numerical calculation (laminar flow and turbulent flow), specifically:
firstly, obtaining heat flow measurement values of all sensors on the surface of the aircraft according to test data, and obtaining heat flow distribution on the surface of the aircraft through statistics; then, calculating full laminar flow and turbulent flow heat flow values of corresponding states at the positions of the sensors by adopting software or a program; finally, comparing the heat flow value identified by the test data with the laminar flow and turbulent flow values obtained by calculation to obtain transition positions of corresponding states at the positions of the sensors;
the software or the program can adopt the existing commercial software, for example, the software can adopt engineering calculation software Fluent to calculate;
for example, referring to FIG. 2, under operating conditions Ma4, Re5 × 107Next, according to the measured values of the heat flows of the sensors on the surface of the aircraft according to the test data, the distribution of the heat flows along the X (flow direction) at the measuring point at the same spanwise position (z-z 0) is obtained; then, calculating laminar flow and turbulent flow heat values corresponding to the test state by adopting engineering calculation software Fluent to obtain the distribution of heat flow along the x direction at the position where z is equal to z 0; finally, comparing the heat flow test value with a turbulence and laminar flow calculation heat flow value to obtain a transition position where z is equal to z0, wherein the heat flow test value deviates from the laminar flow calculation value to a position x0 where the turbulence calculation value develops, namely the transition position, and connecting the transition positions of all z positions to obtain the transition position of the surface of the aircraft;
step two: transition criterion N is obtained by adopting improved e-N stability analysis and judgment method0；
Step 2.1, a neutral curve is obtained by adopting a stability theory;
step 2.1.1, calculating the laminar flow field corresponding to the test state in the step 1.1 by adopting numerical values, and taking the wall temperature in the test state as the wall temperature;
in this step, the numerical calculation is a technique well known in the art, and is described in detail in the literature, "computational fluid dynamics methods and applications", tokyo aeronautics and astronautics university press;
step 2.1.2, the laminar flow field obtained in the step 2.1.1 is converted to an orthogonal grid on the surface of the aircraft along the normal direction, so that an orthogonal laminar flow field is obtained;
step 2.1.3, taking the orthogonalized flow field obtained in the step 2.1.2 as a basic flow field, analyzing the characteristics of a neutral curve by adopting a linear stability theory to obtain a neutral curve considering cross flow, and taking a first mode which is most affected by the cross flow in the neutral curve as a cross flow mode;
in the step, the neutral curve characteristics are analyzed by adopting a linear stability theory by adopting a means known in the field, specifically, see flow stability, Zhongchang, national defense industry publishing company;
for example, referring to fig. 3, in general, when a neutral curve is analyzed by using a linear stability theory, only components in x and y directions are considered in a basic flow, and a component in a z direction is forced to be 0; however, after considering cross-flow, the component in the z-direction in the base stream is no longer forced to be 0, and fig. 3 gives a comparison of the neutral curves when cross-flow is not considered and when cross-flow is considered: in the diagram, the abscissa is the spanwise wave number beta of the disturbance wave, the ordinate is the frequency f of the disturbance wave, and the change of the cloud diagram color represents the amplitude growth rate-alpha of the disturbance waveiOf a size of-alphaiThe set of points for the 0 condition is called the neutral curve. As can be seen from the figure, the low frequency perturbation mode is greatly affected by cross flow, which causes the first mode destabilization region to expand and move to a lower frequency region, compared to the second mode (the Mack mode). The first mode most affected by the cross flow at this time may be regarded as a cross flow mode;
step 2.2 transition criterion N0A method of determining a value;
step 2.2.1 selecting amplitude growth rate-alpha according to the neutral curve obtained in step 2.1.3i>Integrating the second mode disturbance wave of 0 and the transverse flow mode wave of lower frequency along the direction of the potential flow by adopting an e-N method to obtain N value distribution of the disturbance waves of different frequencies along the direction of the potential flow, as shown in FIG. 4;
in this step, the specific calculation formula of the N value is a conventional formula in the e-N method, as follows:
where A is the amplitude of the disturbance, A0 is the amplitude when the disturbance starts to grow, and is generally 0.0001, and x0 represents the disturbance wave at a certain frequency such that the growth rate α isiX denotes a flow direction position of 0;
step 2.2.2, according to the N value distribution of the disturbance waves with different frequencies obtained in step 2.2.1 along the direction of the potential flow, counting to obtain an N value envelope curve (a set of maximum N values at different x positions) along the direction of the potential flow, and forming the N value distribution of the aircraft surface corresponding to the test state in step 1.1, as shown in FIG. 5;
step 2.2.3, based on the result of step 2.2.2, comparing the result with the aircraft surface transition position obtained in the step one, and determining the value N of the trigger transition0As transition criterion of the e-N method;
referring to FIG. 6, at the transition position, where the value of N on the aircraft surface is about 20, N of the transition is triggered020, and the existing empirically derived N0With a great difference of 10, we can see that the prior empirically obtained N0The value is no longer suitable for judging the transition of the boundary layer of the hypersonic aircraft with the complex shape of the lifting body/waverider;
further, said N0The values may be continually corrected based on flight/ground test data.
Step three: the obtained transition criterion N0The method is used for judging the surface transition position of the target aircraft, and specifically comprises the following steps:
step 3.1 ballistic analysis;
aiming at the flight trajectory of a target aircraft, the variation ranges of Mach number, altitude, attack angle and wall surface temperature are obtained by analyzing and counting the variation range of each parameter of the trajectory, and a sample required for establishing a database is obtained, wherein the step is a step known in the field;
step 3.2, establishing an N value distribution database, which comprises the following steps:
step 3.2.1, obtaining a laminar flow field of an aircraft sample working condition;
according to the sample working condition obtained in the step 3.1, adopting engineering calculation software Fluent to calculate and obtain a laminar flow field of the sample working condition of the aircraft, and in the calculation process, on the basis of the aerodynamic thermal environment under the flight trajectory condition and the real aircraft structure, obtaining the temperature of the surface of the aircraft through structural heat transfer analysis, and taking the temperature as the wall surface condition calculated by Fluent;
step 3.2.2, the laminar flow field obtained in the step 3.2.1 is converted to an orthogonal grid on the surface of the aircraft along the normal direction, so that an orthogonal laminar flow field is obtained;
step 3.2.3, taking the orthogonalized flow field obtained in the step 3.2.2 as a basic flow, and obtaining the N value distribution on the surface of the target aircraft, wherein the specific implementation mode comprises the step 2.1.3 and the step 2.2.2;
step 3.3, judging the transition position of the target aircraft;
according to the N-value distribution database established in the step 3.2, for any one ballistic point, a transition criterion N can be established0The transition location of the aircraft surface is determined 20.
The invention is verified by a flight test of a target aircraft according to a new transition criterion N0The target aircraft is predicted by adopting an improved e-N method considering the influence of the cross flow, a transition prediction result of a part with strong three-dimensional property is better matched with a flight test result, and the effectiveness of the transition criterion is explained.
Features that are described and/or illustrated above with respect to one embodiment may be used in the same way or in a similar way in one or more other embodiments and/or in combination with or instead of the features of the other embodiments.
It should be emphasized that the term "comprises/comprising" when used herein, is taken to specify the presence of stated features, integers, steps or components but does not preclude the presence or addition of one or more other features, integers, steps, components or groups thereof.
The above devices and methods of the present invention can be implemented by hardware, or can be implemented by hardware and software. The present invention relates to a computer-readable program which, when executed by a logic section, enables the logic section to realize the above-described apparatus or constituent section, or to realize the above-described various methods or steps. The present invention also relates to a storage medium such as a hard disk, a magnetic disk, an optical disk, a DVD, a flash memory, or the like, for storing the above program.
The many features and advantages of these embodiments are apparent from the detailed specification, and thus, it is intended by the appended claims to cover all such features and advantages of these embodiments which fall within the true spirit and scope thereof. Further, since numerous modifications and changes will readily occur to those skilled in the art, it is not desired to limit the embodiments of the invention to the exact construction and operation illustrated and described, and accordingly, all suitable modifications and equivalents may be resorted to, falling within the scope thereof.
The invention has not been described in detail and is in part known to those of skill in the art.
Claims (5)
1. A judgment method for transition of a boundary layer of a hypersonic aircraft is characterized by comprising the following steps:
step 1, determining a transition position of the surface of the early-stage aircraft;
1.1, acquiring early-stage test data of the aircraft, wherein a required test state is selected according to requirements;
1.2 determining the transition position of the surface of the aircraft in the test state according to the test data;
step 2, obtaining transition criterion N by adopting an improved e-N stability analysis and judgment method0；
2.1 calculating the laminar flow field in the test state, and interpolating the flow field to a grid which is normal and orthogonal to the surface of the aircraft to obtain an orthogonal flow field;
2.2 analyzing a neutral curve by taking the orthogonalized flow field as a basic flow based on the orthogonalized flow field to obtain a neutral curve considering the influence of the transverse flow, wherein a first mode which is most influenced by the transverse flow in the neutral curve is taken as a transverse flow mode;
2.3 selecting a second mode disturbance wave and a cross flow mode disturbance wave with amplitude growth rate larger than 0 based on the obtained neutral curve, integrating the disturbance waves along the direction of potential flow by adopting an e-N method, and taking the envelope curve of the calculated N values of all the disturbance waves as the N value distribution of the surface of the aircraft;
2.4 based on the transition position of the aircraft surface and the N-value distribution of the aircraft surface, determine the N-value transition criterion N for triggering transition0；
Step 3, enabling the obtained transition criterion N0And judging the surface transition position of the target aircraft.
2. The method for judging the transition of the boundary layer of the hypersonic aircraft according to claim 1, characterized in that: in the step 2.1, when the laminar flow field is calculated, the required wall surface temperature is calculated according to the aerodynamic heat along the trajectory and the aircraft structure.
3. The method for judging the transition of the boundary layer of the hypersonic aircraft according to claim 1, characterized in that: said N0The values may be continually corrected based on flight/ground test data.
4. The method for judging the transition of the boundary layer of the hypersonic aircraft according to claim 1, characterized in that: the transition criterion N is obtained0A judgement for a position of surface transition of a target aircraft includes:
obtaining a laminar flow field of a typical working condition of the flight condition of the target aircraft, and interpolating the laminar flow field to a grid which is normal and orthogonal to the surface of the target aircraft to obtain an orthogonal flow field;
based on the orthogonalized flow field, obtaining the N value distribution of the surface of the target aircraft by adopting the same analysis and calculation method of 2.2-2.3 in the step 2;
based on the N value distribution of the target aircraft surface and the N mentioned above0And determining the transition position of the surface of the target aircraft.
5. The method for judging the transition of the boundary layer of the hypersonic flight vehicle according to any one of claims 1 to 4, wherein the hypersonic flight vehicle is a hypersonic flight vehicle with a lifting body/waverider appearance.
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