CN113505542B - Method for constructing turbulence transition model of backswept wing of hypersonic aircraft - Google Patents
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Abstract
The invention discloses a method for constructing a turbulence transition model of a backswept wing of a hypersonic aircraft, which enables the established turbulence transition model to have a Mack mode and a transverse flow disturbance mode simultaneously. A more reasonable model is provided for researching the hypersonic aircraft turbulence transition mechanism and control thereof, so that quantitative analysis can be carried out on the hypersonic aircraft turbulence transition mechanism.
Description
Technical Field
The invention relates to the technical field of aircraft flow field analysis, in particular to a method for constructing a turbulence transition model of a backswept wing of a hypersonic aircraft.
Background
The appearance of the hypersonic aerocraft is mainly a lifting body or a conical shape. Such as the HIFiRe5 standard mode in the united states, or the high angle of attack conic standard mode. The transition from laminar flow to turbulent flow has great influence on the friction resistance and the thermal load of the hypersonic aircraft. For hypersonic aircraft, after the transition occurs, the aircraft drag and surface heat transfer increase dramatically.
For hypersonic aircraft flow, it is typically a three-dimensional boundary layer flow, such as a hypersonic conic flow, a hypersonic lift body profile. For a hypersonic three-dimensional boundary layer, the boundary layer flow can excite a Mack mode and a cross flow disturbance mode. At present, wind tunnel experiments and numerical calculation show that a boundary layer of the wind tunnel is dominated by a cross flow disturbance mode and a Mack mode, and a turbulence transition mechanism of the wind tunnel is also determined by interaction of the cross flow disturbance mode and the Mack mode. The interaction of a transverse flow disturbance mode and a Mack mode is researched by adopting a wind tunnel experiment, and due to the limitation of a measurement condition under a hypersonic speed condition, the pressure pulsation of wall surface and space distribution is mainly measured at present, and whether the measured pressure pulsation is the Mack mode or the transverse flow disturbance mode cannot be distinguished. Therefore, the interaction mechanism of the two modes cannot be well studied in wind tunnel experiments. In order to study the interaction mechanism, direct numerical simulation technology is generally adopted. However, for direct numerical simulation, simple shape and self-matching of boundary conditions are required. Therefore, the contour of the lifting body and the contour of the cone with the attack angle are not suitable for being directly used for numerical simulation of the hypersonic aircraft. Firstly, it is not easy to generate high-precision grids because of the complex shape, and secondly, hypersonic aircrafts such as lifting bodies and cones with large attack angles make the flow in the spanwise direction aperiodic, and cannot be mathematically analyzed by strict mathematical conditions. Therefore, the research on the transition mechanism of the three-dimensional boundary layer turbulence of the hypersonic aircraft is mainly qualitative and not quantitative. Therefore, in order to accurately understand and research the transition mechanism of the hypersonic three-dimensional boundary layer turbulent flow, a more simplified simulation shape needs to be constructed to research the transition mechanism of the hypersonic three-dimensional boundary layer turbulent flow.
Disclosure of Invention
In order to overcome the defects of the prior art, the invention provides a method for constructing a turbulence transition model of a backswept wing of a hypersonic aircraft, and provides a more reasonable model for researching the turbulence transition mechanism and control of the hypersonic aircraft, so that quantitative analysis can be carried out on the turbulence transition model.
The technical scheme adopted by the invention for solving the problems is as follows:
a method for constructing a turbulence transition model of a backswept wing of a hypersonic aircraft enables the established turbulence transition model to have a Mack mode and a cross flow disturbance mode at the same time.
Meanwhile, the method has a Mack mode and a transverse flow disturbance mode, and is convenient for researching interaction mechanisms of the two modes, so that the simulation of the swept-back wing of the hypersonic aircraft is more accurate, the interaction of the Mack mode and the transverse flow disturbance mode in the boundary layer flow of the hypersonic aircraft under the hypersonic condition is convenient to research, the mechanical mechanism of the swept-back wing of the hypersonic aircraft is convenient to analyze more accurately, and the aircraft which is more suitable for real flow field flight is convenient to research, improve and manufacture in the later period.
As a preferable technical scheme, the method comprises the following steps:
s1, establishing the appearance of a swept-back wing of a hypersonic aircraft by adopting a parabolic curve;
s2, generating a numerical calculation grid on the basis of the appearance of the sweepback wing;
s3, calculating a laminar flow field by using computational flow mechanics under the condition of obtaining a computational grid;
s4, after accurate laminar flow is obtained, analyzing whether an unstable Mack mode and a transverse flow disturbance mode exist in a laminar flow field by adopting linear stability;
and S5, confirming that the profile of the sweepback wing can simultaneously have a Mack mode and a crossflow disturbance mode through a linear stability calculation result.
Meanwhile, the method has a Mack mode and a transverse flow disturbance mode, and is convenient for researching interaction mechanisms of the two modes, so that the simulation of the swept-back wing of the hypersonic aircraft is more accurate, the interaction of the Mack mode and the transverse flow disturbance mode in the boundary layer flow of the hypersonic aircraft under the hypersonic condition is convenient to research, the mechanical mechanism of the swept-back wing of the hypersonic aircraft is convenient to analyze more accurately, and the aircraft which is more suitable for real flow field flight is convenient to research, improve and manufacture in the later period.
Specifically, firstly, aiming at the working condition of the hypersonic speed aircraft, the working condition is consistent with the incoming flow of the hypersonic speed aircraft. And determining the Mach number and the wall surface temperature of the incoming flow. And simulating a local flow field of the hypersonic aircraft by adopting the hypersonic three-dimensional sweepback wings, and determining the sweepback angle and the stagnation point radius of the hypersonic three-dimensional sweepback wings. And calculating the hypersonic three-dimensional sweepback wing laminar flow field by adopting a high-precision numerical method. After obtaining the accurate flow field, calculating the N values of the Mack mode and the transverse flow disturbance mode by adopting a linear stability theory, and after confirming that the hypersonic three-dimensional sweepback wing laminar flow field has two modes together, carrying out disturbance interaction analysis by utilizing the flow field.
As a preferable technical scheme, the method comprises the following steps: setting a spanwise distribution of the turbulence transition model to have a periodic boundary condition.
The flow field is periodic in the spanwise direction, the defect that a transition mechanism of the hypersonic aircraft cannot be well researched under the wind tunnel experiment condition is overcome, and a series of analysis can be conveniently carried out on the flow field.
As a preferred technical scheme, the method comprises the following steps: setting a spanwise direction of the turbulence transition model to infinite spanwise flow.
Spanwise infinite spanwise flow is a good periodic flow, and the flow field mechanism is more convenient to analyze.
As a preferred solution, in step S1, the created longitudinal section of the profile of the swept-back wing of the aircraft conforms to the following functional image shape:
(x,y)|y 2 ＝kRx；
wherein x is the chord length of the airfoil, and the unit is m; r is the radius of the leading edge of the swept-back wing, and the unit is m; k is an adjusting coefficient, is a normal number and has no unit; y is the vertical chord length direction.
The shape and the appearance of the function image are simple, and a high-precision numerical grid can be generated; secondly, the flow field is periodic in the spanwise direction so as to be convenient for carrying out a series of analysis on the flow field; meanwhile, the camber design can design the sweepback wing according to the radius R of the leading edge of the sweepback wing, the flow of the simplified camber under the hypersonic speed condition simultaneously comprises a Mack mode and a transverse flow disturbance mode, the mechanism of turbulent flow transition induced by the interaction of the Mack mode and the transverse flow disturbance mode in the boundary layer flow of the aircraft under the hypersonic speed condition can be conveniently researched, and the mechanical mechanism of the sweepback wing of the hypersonic speed aircraft can be conveniently and accurately analyzed. Especially in the case of an infinite span wing with a large sweep angle (e.g., 65 degrees).
As a preferred technical scheme, the angle range of the leading edge sweepback angle of the sweepback wing is selected to be 55-70 degrees.
As a preferred technical scheme, the sweep angle range of the trailing edge of the swept wing is selected to be 55-70 degrees.
The arrangement of the front edge sweepback angle of the sweepback wing is more consistent with the actual flow field scene of interaction of a Mack mode and a transverse flow disturbance mode, and the simulation is more real. Preferably, the trailing edge sweep angle of the swept wing is equal to the leading edge sweep angle of the swept wing.
As a preferable technical scheme, the critical current Mach number of the swept-back wing is 5-7.
The setting of the critical incoming flow Mach number of the sweepback wing is more in line with the actual flow field scene of the interaction of the Mack mode and the cross flow disturbance mode, and the Mach number interval calculation can reflect the actual flow field more and the simulation is more real.
As a preferred technical scheme, the range of k is selected to be 0-1.
Compared with the prior art, the invention has the following beneficial effects:
(1) The simulation appearance can grasp the main characteristic of turbulent flow transition of the hypersonic aircraft flow field, namely a Mack mode and a transverse flow disturbance mode coexist in the flow field, the spanwise direction is infinite wingspan flow, the flow is periodic flow, and a strict mathematical theory analysis flow field mechanism can be adopted; the method is convenient for researching a turbulent flow transition mechanism induced by interaction of a Mack mode and a transverse flow disturbance mode under the condition of high supersonic velocity, and is convenient for analyzing the turbulent flow transition mechanism, so that the aircraft which is more suitable for actual flow field flight is developed and improved in the later period;
(2) The arrangement of the leading edge sweepback angle of the sweepback wing and the critical incoming flow Mach number of the sweepback wing better conforms to the actual flow field scene of the interaction of a Mack mode and a cross flow disturbance mode.
Drawings
FIG. 1 is a hypersonic swept-back airfoil profile and pressure profile under high-velocity flow; wherein, the abscissa represents the chord length, and the unit is m; the ordinate represents the vertical chord length direction, and the unit is m;
FIG. 2 is a hypersonic swept-back airfoil profile;
FIG. 3 is a steady cross flow disturbance mode in laminar flow of a boundary layer of a swept-back wing with a hypersonic infinite span; wherein, the abscissa represents the chord length direction, and the unit is m; the ordinate represents the integral of the growth rate of the disturbance mode along the chord length, and the integral of the growth rate of the disturbance mode greater than 0 represents that the mode is unstable and is a dimensionless parameter;
FIG. 4 is a Mack mode in laminar flow of a boundary layer of a hypersonic infinite span sweepback wing; wherein, the abscissa represents the chord length direction, and the unit is m; the ordinate represents the integral of the growth rate of the disturbance mode along the chord length, and the condition that the mode is unstable and is a dimensionless parameter is represented by being larger than 0.
Detailed Description
The present invention will be described in further detail with reference to examples and drawings, but the present invention is not limited to these examples.
Examples
As shown in fig. 1 to 4, a method for constructing a turbulence transition model of a backswept wing of a hypersonic aircraft enables the established turbulence transition model to have a Mack mode and a cross flow disturbance mode at the same time.
Meanwhile, the method has a Mack mode and a transverse flow disturbance mode, and is convenient for researching interaction mechanisms of the two modes, so that the simulation of the swept-back wing of the hypersonic aircraft is more accurate, the interaction of the Mack mode and the transverse flow disturbance mode in the boundary layer flow of the hypersonic aircraft under the hypersonic condition is convenient to research, the mechanical mechanism of the swept-back wing of the hypersonic aircraft is convenient to analyze more accurately, and the aircraft which is more suitable for real flow field flight is convenient to research, improve and manufacture in the later period.
As a preferable technical scheme, the method comprises the following steps:
s1, establishing the appearance of a swept-back wing of a hypersonic aircraft by adopting a parabolic curve;
s2, generating a numerical calculation grid on the basis of the appearance of the sweepback wing;
s3, calculating a laminar flow field by using computational flow mechanics under the condition of obtaining a computational grid;
s4, after accurate laminar flow is obtained, analyzing whether an unstable Mack mode and a transverse flow disturbance mode exist in a laminar flow field by adopting linear stability;
and S5, confirming that the sweepback wing profile can simultaneously have a Mack mode and a cross flow disturbance mode through a linear stability calculation result.
Meanwhile, the method has a Mack mode and a transverse flow disturbance mode, and is convenient for researching interaction mechanisms of the two modes, so that the simulation of the swept-back wing of the hypersonic aircraft is more accurate, the interaction of the Mack mode and the transverse flow disturbance mode in the boundary layer flow of the hypersonic aircraft under the hypersonic condition is convenient to research, the mechanical mechanism of the swept-back wing of the hypersonic aircraft is convenient to analyze more accurately, and the aircraft which is more suitable for real flow field flight is convenient to research, improve and manufacture in the later period.
Specifically, the working condition of the hypersonic flight vehicle is firstly consistent with the incoming flow of the hypersonic flight vehicle. And determining the Mach number and the wall surface temperature of the incoming flow. And simulating a local flow field of the hypersonic aircraft by adopting the hypersonic three-dimensional sweepback wings, and determining the sweepback angle and the stagnation point radius of the hypersonic three-dimensional sweepback wings. And calculating a hypersonic three-dimensional sweepback wing laminar flow field by adopting a high-precision numerical method. After obtaining the accurate flow field, calculating the N values of the Mack mode and the transverse flow disturbance mode by adopting a linear stability theory, and after confirming that the hypersonic three-dimensional sweepback wing laminar flow field has two modes together, carrying out disturbance interaction analysis by utilizing the flow field.
As a preferred technical scheme, the method comprises the following steps: setting a spanwise distribution of the turbulence transition model to have a periodic boundary condition.
The flow field is periodic in the spanwise direction, the defect that a transition mechanism of the hypersonic aircraft cannot be well researched under the wind tunnel experiment condition is overcome, and a series of analysis can be conveniently carried out on the flow field.
As a preferable technical scheme, the method comprises the following steps: setting a spanwise direction of the turbulence transition model to infinite spanwise flow.
Spanwise infinite spanwise flow is a good periodic flow, and the flow field mechanism is more convenient to analyze.
As a preferred solution, in step S1, the created longitudinal section of the profile of the swept-back wing of the aircraft conforms to the following functional image shape:
(x,y)|y 2 ＝kRx；
wherein x is the chord length of the airfoil and the unit is m; r is the radius of the leading edge of the swept-back wing, and the unit is m; k is an adjusting coefficient, is a normal number and has no unit; y is the vertical chord length direction.
The shape and the appearance of the function image are simple, and a high-precision numerical grid can be generated; secondly, the flow field is periodic in the spanwise direction so as to be convenient for carrying out a series of analysis on the flow field; meanwhile, the camber design can design the sweepback wing according to the radius R of the leading edge of the sweepback wing, the flow of the simplified camber under the hypersonic speed condition simultaneously comprises a Mack mode and a transverse flow disturbance mode, the mechanism of turbulent flow transition induced by the interaction of the Mack mode and the transverse flow disturbance mode in the boundary layer flow of the aircraft under the hypersonic speed condition can be conveniently researched, and the mechanical mechanism of the sweepback wing of the hypersonic speed aircraft can be conveniently and accurately analyzed. Especially in the case of an infinite span wing with a large sweep angle (e.g., 65 degrees).
The linear stability of the laminar flow field of the hypersonic three-dimensional boundary layer under the hypersonic Mach6 condition is calculated, so that the hypersonic infinite span sweepback wing designed by the invention really has a Mack mode and a cross flow disturbance mode together under the hypersonic condition, as shown in fig. 3 and 4. FIG. 3 shows the steady transverse flow vortex mode N value in the laminar boundary layer flow of the infinite span sweepback wing under the hypersonic speed condition. The value of N indicates that the unstable disturbance mode is integrated along the flow direction, so that the disturbance is unstable in the present case as long as the value of N is greater than 0. Therefore, FIG. 3 shows that the simplified outline flow field has unstable steady cross-flow disturbance mode with the spanwise wavelength of 20mm-40 mm. While figure 4 shows the N value of the Mack mode at different frequencies with an spanwise wavenumber of 16mm, this indicates that the unstable Mack mode exists throughout the flow field for the simplified profile.
As a preferred technical scheme, the angle range of the leading edge sweepback angle of the sweepback wing is selected to be 55-70 degrees.
As a preferred technical scheme, the sweep angle range of the trailing edge of the swept wing is selected to be 55-70 degrees.
The arrangement of the front edge sweepback angle of the sweepback wing is more consistent with the actual flow field scene of interaction of a Mack mode and a transverse flow disturbance mode, and the simulation is more real. Preferably, the trailing edge sweep angle of the swept wing is equal to the leading edge sweep angle of the swept wing.
As a preferable technical scheme, the critical current Mach number of the swept-back wing is 5-7.
The setting of the critical incoming flow Mach number of the sweepback wing is more consistent with the actual flow field scene of interaction of a Mack mode and a cross flow disturbance mode, and the Mach number interval calculation can reflect a real flow field more and the simulation is more real.
As a preferred technical scheme, the range of k is selected to be 0-1.
As described above, the present invention can be preferably realized.
The foregoing is only a preferred embodiment of the present invention, and the present invention is not limited thereto in any way, and any simple modification, equivalent replacement and improvement made to the above embodiment within the spirit and principle of the present invention still fall within the protection scope of the present invention.
Claims (5)
1. A method for constructing a turbulence transition model of a backswept wing of a hypersonic aircraft is characterized in that the established turbulence transition model has a Mack mode and a transverse flow disturbance mode simultaneously;
the method comprises the following steps:
s1, establishing the appearance of a swept-back wing of a hypersonic aircraft by adopting a parabolic curve;
s2, generating a numerical calculation grid on the basis of the appearance of the sweepback wing;
s3, under the condition of obtaining a computational grid, calculating a laminar flow field by using computational flow mechanics;
s4, after accurate laminar flow is obtained, analyzing whether an unstable Mack mode and a transverse flow disturbance mode exist in a laminar flow field by adopting linear stability;
s5, confirming that the profile of the sweepback wing can have a Mack mode and a cross flow disturbance mode simultaneously through a linear stability calculation result;
further comprising the steps of: setting a spanwise of the turbulence transition model to have a periodic boundary condition;
further comprising the steps of: setting a spanwise distribution of the turbulence transition model to an infinite spanwise flow;
in step S1, the established longitudinal section of the profile of the swept-back wing of the aircraft conforms to the following functional image shape:
(x,y)|y 2 ＝kRx；
wherein x is the chord length of the airfoil and the unit is m; r is the radius of the leading edge of the swept-back wing, and the unit is m; k is an adjusting coefficient, is a normal number and has no unit; and y is the vertical chord length direction.
2. The method for constructing the turbulence transition model of the swept-back wings of the hypersonic aircraft according to claim 1, wherein the range of the leading edge swept-back angle of the swept-back wings is selected to be 55-70 °.
3. The method for constructing the turbulence transition model of the swept-back wings of the hypersonic aircraft according to claim 1, wherein the sweep angle range of the trailing edges of the swept-back wings is selected to be 55-70 °.
4. The method for constructing the turbulence transition model of the swept-back wing of the hypersonic aircraft according to claim 1, wherein the critical incoming flow Mach number of the swept-back wing is 5-7.
5. The method for constructing the turbulence transition model of the backswept wing of the hypersonic aircraft according to claim 1, wherein k is selected to be in a range of 0-1.
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| CN108197388A (en) * | 2018-01-02 | 2018-06-22 | 清华大学 | A kind of acquisition methods and system of high-speed aircraft flow transition characteristic |
| CN110481761A (en) * | 2019-08-20 | 2019-11-22 | 空气动力学国家重点实验室 | It is a kind of to utilize surface aperture/slot flow transition passive control device |
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| CN108304601B (en) * | 2017-08-09 | 2019-12-31 | 北京空天技术研究所 | Method for judging transition of boundary layer of hypersonic aircraft |
| CN108304600B (en) * | 2017-08-09 | 2020-03-31 | 北京空天技术研究所 | Method for predicting transition position of hypersonic aircraft |
| CN108009320B (en) * | 2017-11-14 | 2021-07-27 | 南京航空航天大学 | Control-oriented multi-system association modeling method for hypersonic aircraft |
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| CN111651897B (en) * | 2020-06-18 | 2022-10-11 | 北京理工大学 | Large-size gap microstructure capable of inhibiting transition of hypersonic velocity boundary layer |
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| CN108197388A (en) * | 2018-01-02 | 2018-06-22 | 清华大学 | A kind of acquisition methods and system of high-speed aircraft flow transition characteristic |
| CN110481761A (en) * | 2019-08-20 | 2019-11-22 | 空气动力学国家重点实验室 | It is a kind of to utilize surface aperture/slot flow transition passive control device |
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