CN116104647B - Air inlet channel design method based on frequency-adjustable oscillation type Ramp type vortex generator - Google Patents

Air inlet channel design method based on frequency-adjustable oscillation type Ramp type vortex generator Download PDF

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CN116104647B
CN116104647B CN202211121374.XA CN202211121374A CN116104647B CN 116104647 B CN116104647 B CN 116104647B CN 202211121374 A CN202211121374 A CN 202211121374A CN 116104647 B CN116104647 B CN 116104647B
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vortex generator
air inlet
frequency
inlet channel
shock wave
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CN116104647A (en
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张悦
王梦格
薛洪超
何小明
谭慧俊
汪昆
陈亮
郑飞飞
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Nanjing University of Aeronautics and Astronautics
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Nanjing University of Aeronautics and Astronautics
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Other Liquid Machine Or Engine Such As Wave Power Use (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

The invention discloses an air inlet channel design method of an oscillating type Ramp type vortex generator based on adjustable amplitude frequency, which is divided into two parts, namely design and installation position determination of the oscillating type Ramp type vortex generator with adjustable frequency, wherein the two parts are used for analyzing interference characteristics of a shock wave boundary layer of an air inlet channel lip cover. Judging whether boundary layer separation is caused under different incoming flow Mach numbers according to the influence mechanism of boundary layer interference of the lip cover shock wave of the air inlet channel and an empirical formula and by the magnitude of static pressure before and after the incidence of the lip cover shock wave; according to the control mechanism of the frequency-adjustable oscillating Ramp type vortex generator on the disturbance of the shock wave boundary layer, the selection criteria of the size and the installation position of the vortex generator are given. Therefore, a convenient and effective air inlet control measure design method is obtained.

Description

Air inlet channel design method based on frequency-adjustable oscillation type Ramp type vortex generator
Technical Field
The invention belongs to the field of supersonic air inlet channel design, and particularly relates to an air inlet channel design method based on a frequency-adjustable oscillating Ramp type vortex generator.
Background
With the rapid development of aerospace technology, the flying speed of aircrafts is also increasing. In recent years, in order to meet the demands of future aerospace integrated flight, hypersonic civil aviation transportation and the like, a hypersonic aircraft with Mach numbers of 0-6+ and capable of taking off and landing horizontally and a combined propulsion technology thereof have become one of the high points in the current international aerospace field. For an air-breathing aircraft, the air intake duct is subject to the heavy duty of providing high quality airflow to the engine and therefore has a critical effect on the overall engine and hence the aircraft. In hypersonic aircraft using combined power in particular, the inlet duct is often of adjustable form in order to accommodate extremely wide flight speed ranges. The adjustable air inlet passage needs to face an extremely wide speed domain and an airspace (Mach numbers 0-6+ and heights 0-30 km) of an upstream flying incoming flow, and a plurality of working flow passages and working modes (the number of the working flow passages is more than or equal to 2 and the number of the working modes is more than or equal to 3) of a downstream engine, and meanwhile, the heavy duty of the combined power mode conversion is also borne, so that the complexity of the air inlet passage is far more than that of an air inlet passage of a traditional aircraft (Mach numbers 0-2 and heights 0-20 km; single flow passage and single mode). For this reason, the design parameters, the working performance and the operating capacity of the supersonic adjustable air intake channel have a decisive influence on the geometry, the working efficiency and the working envelope of the supersonic aircraft, which significantly exceeds the conventional air intake channels of aircraft.
The capturing and the decelerating pressurization of the incoming flow of the air inlet channel are mainly realized by means of a shock wave system, and the boundary layer on the internal compression surface of the air inlet channel is continuously developed when the air inlet channel is sucked into the boundary layer of the upstream machine body, so that the shock wave boundary layer interference becomes an unavoidable flow phenomenon in the air inlet channel. Considering the requirement of reducing the external resistance of the air inlet channel, the included angle between the inner surface of the lip cover of the air inlet channel and the free incoming flow can not be too large, which means that the air flow after the precursor is compressed can face a larger geometric compression angle at the lip cover, the reverse pressure gradient applied by the lip cover shock wave induced at the position to the boundary layer at the compression surface side often exceeds the separation limit of the local boundary layer, and at the moment, if the design is improper, obvious separation can occur at the air inlet port part, and the working state of the air inlet channel is further deviated from the design working condition. Under certain extreme conditions, shock-induced large-scale separation can also block the flow channel, so that the air inlet channel is not started, and the working state of the aircraft is seriously influenced. Therefore, the performance of the supersonic air intake is most significantly affected by the incident shock/boundary layer disturbance phenomenon represented by the lip shroud shock/boundary layer disturbance. In addition, such disturbances tend to occur upstream of the channels within the inlet, and complex additional wave trains and unsteady disturbances induced by local shock/boundary layer disturbances may propagate downstream and affect the overall downstream channel flow field. In summary, the incident shock/boundary layer interference phenomenon is a fundamental problem in supersonic air inlets.
The design method of the air inlet channel aiming at the problem comprises the following several common schemes: the lip cover changing scheme designs the internal pressure section, the throat and the combustion chamber of the air inlet channel as a whole, which is difficult to realize in practical application; the boundary layer gassing enhances the ability to resist the reverse pressure gradient by removing low energy fluid within the boundary layer, thereby achieving the purpose of inhibiting separation. However, the boundary layer deflation method can obtain a control effect and simultaneously inevitably needs to deflate a part of air flow captured by the air inlet channel; the height of the miniature slope type vortex generator is only 10% -70% of the thickness of the boundary layer, so that the control effect is achieved, the additional resistance of the control part is greatly reduced, and the miniature slope type vortex generator attracts attention of researchers.
The main advantage of vortex generators over other flow control techniques is that no additional energy or mass supply is required, the generated flow vortices can be kept a long distance in the boundary layer and are not easily lifted off the boundary layer. However, from the existing research results, the miniature slope type vortex generator can only effectively control shock waves/boundary layer interference under certain conditions, however, the flow field in the actual supersonic speed air inlet channel is very complex, and the control capability of the fixed geometry vortex generator can not reach the actual control requirement of the air inlet channel. Obviously, by combining the flow characteristics of the air inlet channel, a new type of micro vortex generator is further developed, and a new control strategy and a control method are provided to realize better control of the interference of the incident shock wave/boundary layer in the air inlet channel.
Disclosure of Invention
In order to solve the problems, the invention provides an air inlet channel design method based on a frequency-adjustable oscillating Ramp type vortex generator. Judging whether boundary layer separation is caused under different incoming flow Mach numbers according to the wave system structure of the designed air inlet channel and an empirical formula; according to the control mechanism of the frequency-adjustable oscillating Ramp type vortex generator on the disturbance of a shock wave boundary layer, the selection criteria of the size and the installation position of the vortex generator are given; according to different flight Mach numbers and flight attack angles of inlet of the air inlet channel, the oscillation type Ramp vortex generator capable of adjusting the frequency can induce high-intensity oscillation flow to the vortex structure through high-frequency oscillation, so that stronger momentum blending effect in the boundary layer is induced, and the control of the interference of the strong shock wave boundary layer is realized.
In order to achieve the above purpose, the air inlet channel based on the frequency-adjustable oscillating Ramp type vortex generator adopts the following technical scheme:
(1) Analysis of boundary layer interference characteristics of shock waves of the lip cover of the air inlet channel:
an air inlet channel based on an oscillation type Ramp type vortex generator capable of adjusting frequency comprises the oscillation type Ramp type vortex generator, an air inlet channel outer compression surface, a lip cover, an air inlet channel inner channel surface, an air inlet channel side wall, an oscillation shaft of the vortex generator, an air inlet channel inner channel, lip cover incident shock waves, reflected shock waves, oscillation type Ramp type vortex generator tail edge points, separation bubble front edge points, air inlet channel lip cover shock wave non-sticking incidence points, separation bubbles and separation bubble tail edge points; the lip cover is positioned outside the surface of the inner channel, and the lip cover and the surface of the inner channel form an inner channel of the air inlet channel; the adjustable frequency oscillation type Ramp vortex generator is placed on the surface of an inner channel of the air inlet channel, a certain distance is reserved on the upstream of a non-sticky incidence point of the lip cover shock wave, and when the influence of separation on the performance of the air inlet channel is small, the vortex generator is embedded into the lower surface of the inner channel to keep the wall surface flat; when the separation has a relatively large influence on the performance of the air inlet channel, the vortex generator can oscillate up and down through a controllable frequency.
The air inlet based on the frequency-adjustable oscillating Ramp type vortex generator is a binary supersonic air inlet, the working Mach number range is 0-3, and the internal contraction ratio of the air inlet is 1.7; the two-stage compression inclined planes are adopted, the wedge angles of the two inclined planes are 9 degrees, and the compression angles of the two-stage lip covers are 11 degrees and 8 degrees respectively; the boundary layer thickness of the inlet section of the air inlet is 1/15 of the inlet height h of the air inlet.
And calculating Mach numbers and pressures of the lip cover after the incidence of the laser and the Mach numbers and pressures of the lip cover after the reflection of the laser according to the oblique laser theory. The oblique shock theory is that the Mach number M 0 When supersonic airflow of the object plane encounters the object plane and the flowing direction deflects by delta (namely, the wedge angle of the object plane), an oblique shock wave is emitted from the starting point of the turning of the object plane, and the included angle between the wave plane and the airflow direction before turning is the oblique shock waveThe wave angle beta, the angle delta of the airflow deflection and the Mach number of the incoming flow can be used for obtaining the angle beta of the oblique shock wave and the Mach number M of the oblique shock wave 1 。
And similarly, according to the intensity and the wave front Mach number of the lip cover incident laser wave, the principle deduced by adopting a free interference theory is expressed as the relation between the pressure after separation and upstream parameters, and whether the inlet lip cover incident laser wave of the inlet channel under different incoming flow Mach numbers can cause obvious boundary layer separation is predicted.
Wherein M is 0 Wave front Mach number, P, of incident shock wave for lip mask 0 For wave front static pressure of incident shock wave, P 1 For the static pressure value after reflection and excitation, C f，0 And gamma is the adiabatic index of air, which is the wall friction coefficient.
(2) Frequency-adjustable oscillating Ramp type vortex generator size and mounting position selection criteria:
the arrangement direction of the vortex generator arranged on the lower surface of the inner channel of the air inlet channel is perpendicular to the supersonic incoming flow direction.
Further, the spacing s=7.5h between the center lines of the frequency-adjustable oscillating Ramp type vortex generator v (wherein h v Height of vortex generator).
Further, the frequency-adjustable oscillating Ramp type vortex generator is arranged between two side walls of the air inlet channel along the spanwise direction.
Further, the basic configuration of the oscillating Ramp type vortex generator is consistent with that of a classical Anderson type slope type vortex generator, the half apex angle AP of the vortex generator is 24 degrees, the bottom of the vortex generator is isosceles triangle, the thickness d of the vortex generator is 0.5mm, and the slope chord length c=7.2 h v The vortex generator oscillates up and down about the oscillation axis at a certain frequency (see fig. 5).
Further, the development of increased momentum flux downstream of the tunable oscillating Ramp vortex generator is measured by the following method. Obtaining an accurate value by calculating a correlation coefficient R between the size of the separation area and the momentum flux E, wherein the correlation coefficient R has an obvious maximum value at a distance of about Y=0.43delta from the wall surface; finally, it can be derived from the following formula that at a height y=0.43 δ (using a height y=0.43 δ because the separation bubble is most sensitive to the momentum flux at a height of 43% δ, i.e. the correlation coefficient vmax), the vortex generator requires a fixed length of x=5.7δ to allow the boundary layer to mix sufficiently to inject the maximum momentum in the near-wall region; delta is the boundary layer thickness;
wherein x is the distance from the downstream to the tail edge of the vortex generator, E (x) is the magnitude of momentum flux at different flow direction positions downstream of the vortex generator, U is the speed of the vortex generator, U clean For speed without vortex generator, U ∞ Is the far forward incoming flow velocity.
Further, from the relationship between E (x) and x in the above formula, it can be seen that: x/δ=5.7 is preceded by a boundary layer intimate mixing region, the high momentum fluid is transported to the surface by the flow direction vortex, and the momentum flux E tends to increase; x/δ=5.7 followed by a plateau region of length 3.8δ, the momentum flux E becomes approximately constant; and the shock impact region evident at x/δ=9.5 (corresponding to a distance of about 7.8 δ upstream from the non-stick incidence location).
Furthermore, the distance between the trailing edge point of the frequency-adjustable oscillating Ramp type vortex generator and the non-sticky incident point of the incident shock wave is larger than L, so that an ideal control effect is achieved.
Where l=7.8δ+5.7δ.
Furthermore, the oscillation height of the frequency-adjustable oscillation type Ramp type vortex generator is 10% -70% of the thickness of the local boundary layer.
Furthermore, the oscillation height and frequency of the frequency-adjustable oscillation type Ramp type vortex generator can be adjusted according to the actual control requirement of the air inlet channel. Controlling the frequency of up-and-down reciprocating oscillation of the vortex generator by controlling the rotating speed of the motor; the appropriate actuating mechanism is selected to achieve the oscillating height of the frequency-tunable oscillating vortex generator.
The beneficial effects are that: the invention provides an air inlet channel based on a frequency-adjustable oscillating Ramp type vortex generator, which can realize the height and frequency change under the drive of an actuating mechanism, thereby inducing a higher-strength frequency-controllable flow direction vortex to realize a stronger momentum mixing effect, breaking through the limitation that the traditional fixed geometry vortex generator can only act on specific working conditions, and greatly expanding the action range of the vortex generator. Because the materials and the configuration of the actuating mechanism are diversified, the actuating mechanism can be designed to generate required vibration height or vibration frequency according to specific working environment and action effect, so that the problem of flow separation of the air inlet under different working conditions can be effectively solved, the flow field quality of the air inlet is improved, the back pressure resistance of the air inlet is improved, and the performance of the whole air inlet is further improved.
Drawings
FIG. 1 is a two-dimensional view of an inlet channel of the present invention based on a frequency tunable oscillating Ramp type vortex generator.
Fig. 2 is a top view of an inlet channel based on a frequency-tunable oscillating Ramp type vortex generator according to the present invention.
Fig. 3 is an initial state of a vortex generator in an air intake based on a frequency-tunable oscillation type Ramp type vortex generator of the present invention.
Fig. 4 shows the oscillation of the vortex generator to the highest state in the inlet channel based on the frequency-adjustable oscillating Ramp type vortex generator according to the present invention.
Fig. 5 is a block diagram of a vortex generator based on an inlet channel of a frequency-tunable oscillating Ramp type vortex generator according to the present invention.
Fig. 6 is a schematic diagram of the boundary layer velocity profile variation of the oscillating Ramp type vortex generator symmetry plane at different moments in time at a cross section x=10hv from the vortex generator trailing edge.
Detailed Description
The invention discloses an air inlet channel based on an oscillation type Ramp type vortex generator capable of adjusting frequency and a design method thereof, and the technical scheme provided by the invention is described in detail below with reference to the accompanying drawings.
1. Analysis of boundary layer interference characteristics of shock waves of the lip cover of the air inlet channel:
(1) As shown in fig. 1, an embodiment of an air inlet channel based on a frequency-adjustable oscillating Ramp type vortex generator is provided. An air inlet channel based on an oscillation type Ramp type vortex generator capable of adjusting frequency comprises the oscillation type Ramp type vortex generator 1, an air inlet channel outer compression surface 2, a lip cover 3, an air inlet channel inner channel surface 4, an air inlet channel side wall 5, a vortex generator oscillation shaft 6 and an air inlet channel inner channel 7; the lip cover 3 is positioned outside the inner channel surface 4, and the lip cover 3 and the inner channel surface 4 form an inner channel 7 of the air inlet channel; the adjustable frequency oscillation type Ramp vortex generator 1 is placed on the inner channel surface 4 of the air inlet channel, a certain distance is reserved on the upstream of a non-sticky incidence point 12 of the lip cover shock wave, and when the influence of separation on the performance of the air inlet channel is small, the vortex generator 1 is embedded into the lower surface 4 of the inner channel to keep the wall surface flat; when the separation has a relatively large influence on the performance of the inlet channel, the vortex generator 1 can oscillate up and down by a controllable frequency, which means a vibratory reciprocating movement between the highest oscillating position and the initial state position. As shown in fig. 3, 4 and 5, 1.1, which is the highest position of the vortex generator 1, the vortex generator 1.2 returns to the initial state, i.e. is in a state of being attached to the inner channel surface 4 (as shown in fig. 3, 4 and 5, position 1.2)
(2) As shown in fig. 1 and fig. 2, fig. 3 and fig. 4, the air inlet is a binary supersonic air inlet, the working Mach number range is 0-3, and the internal contraction ratio of the air inlet is 1.7; the two-stage compression inclined plane 2 is adopted, the wedge angle of the two inclined planes is 9 degrees, and the compression angles of the two-stage lip covers are 11 degrees and 8 degrees respectively; the boundary layer thickness delta of the inlet section of the air inlet channel is 1/15 of the throat height h of the air inlet channel.
(3) As shown in fig. 1, according to the incoming flow parameters and geometric parameters of the air inlet, the flow characteristics of the air inlet channel 4 within a preset working Mach number range are analyzed according to the oblique shock wave theory, and the wave system and the bubble separation structure (see fig. 1) are drawn, including a lip cover incident shock wave 8, a reflected shock wave 9 and an oscillating Ramp type vortex generator tailA rim point 10, a separation bubble leading edge point 11, an inlet lip shroud shock wave non-stick incidence point 12, a separation bubble 13, and a separation bubble trailing edge point 14. The oblique shock theory is that the Mach number M 0 When supersonic airflow of the (a) encounters an object plane and the flow direction deflects by delta (namely, the wedge angle of the object plane), an oblique shock wave is emitted from the starting point of the turning of the object plane, the included angle between the wave plane and the airflow direction before the turning is the oblique shock wave angle beta, and the oblique shock wave angle beta and the Mach number M after the oblique shock wave can be obtained from the airflow deflection angle delta and the incoming flow Mach number 1 . Similarly, the Mach number and pressure of the lip mask after the incident shock wave 8 and the Mach number and pressure of the lip mask after the reflected shock wave 9 are calculated according to the oblique shock wave theory.
(4) And, the criteria derived by the free interference theory is expressed as the relationship between the pressure after separation and the upstream parameter, and whether the incoming shock wave 8 of the inlet lip shroud at different incoming flow Mach numbers can all cause obvious boundary layer separation can be predicted by the parameters.
Wherein M is 0 Wave front Mach number, P, of incoming shock wave 8 for lip shroud 0 For the wavefront static pressure of the incident shock wave 8, P 1 For static pressure value after reflection of shock wave 9, C f，0 And gamma is the adiabatic index of air, which is the wall friction coefficient.
2. Frequency-adjustable oscillating Ramp type vortex generator size and mounting position selection criteria:
(1) As shown in fig. 1, 2 and 5, the arrangement direction of the vortex generators 1 arranged on the lower surface 4 of the inner channel of the inlet of the air inlet is perpendicular to the supersonic incoming flow direction, and the space s=7.5 hv between the center lines of the vortex generators 1 (where hv is the height of the vortex generators) is arranged between the two side walls 5 of the air inlet along the spanwise direction. The basic configuration of the oscillation type Ramp type vortex generator is consistent with that of a classical Anderson type slope type vortex generator, the half apex angle AP of the vortex generator 1 is 24 degrees, the bottom of the vortex generator is an isosceles triangle, the thickness d of the vortex generator is 0.5mm, the chord length c=7.2 hv of the slope, and the vortex generator swings up and down around an oscillation shaft.
(2) The development of increased momentum flux downstream of the tunable oscillating Ramp type vortex generator is measured by the following method. Obtaining an accurate value by calculating a correlation coefficient R between the size of the separation area and the momentum flux E, wherein the correlation coefficient R has an obvious maximum value at a distance of about Y=0.43delta from the wall surface; finally, it can be derived from the following formula that at a height y=0.43 δ (using a height y=0.43 δ because the separation bubble is most sensitive to the momentum flux at a height of 43% δ, i.e. the correlation coefficient vmax), the vortex generator 1 requires a fixed length of x=5.7δ to allow the boundary layer to mix sufficiently to inject the maximum momentum in the near-wall region; delta is the boundary layer thickness;
wherein x is the distance from the downstream to the tail edge of the vortex generator 1, E (x) is the magnitude of momentum flux at different flow direction positions downstream of the vortex generator 1, U is the speed of the vortex generator 1, U clean U is the speed of the vortex generator 1 ∞ Is the far forward incoming flow velocity.
(3) As shown in fig. 1, the position of a non-sticky incident point 12 of the shock wave of the lip cover of the air inlet channel is obtained in the Mach number range required to be controlled by the air inlet channel based on the frequency-adjustable oscillation type Ramp type vortex generator, and the position interval where the interference of the shock wave boundary layer occurs, namely the distance between a separation bubble leading edge point 11 and a separation bubble trailing edge point 14, is determined. Wherein the shock boundary layer disturbance impact zone upstream of the corresponding shock non-stick incident point 12 is 7.8 delta. The distance between the trailing edge point of the vortex generator 1 and the incident shock wave non-stick incident point 12 is greater than L to achieve the desired control effect. Where l=7.8δ+5.7δ.
(4) As shown in FIG. 5, the oscillation height of the frequency-adjustable oscillation type Ramp type vortex generator is 10% -70% of the thickness of the local boundary layer. The oscillation height and frequency of the frequency-adjustable oscillation type Ramp type vortex generator 1 can be adjusted according to the actual control requirement of an air inlet channel. Controlling the frequency of up-and-down reciprocating oscillation of the vortex generator 1 by controlling the rotating speed of the motor; the appropriate actuating mechanism is selected to achieve the oscillation height of the frequency-tunable oscillating vortex generator 1. High-intensity oscillation flow is induced to flow to the vortex structure through high-frequency oscillation, and stronger momentum blending effect in the boundary layer is induced to realize control of interference of a strong shock boundary layer.
Examples
According to the steps, the air inlet channel based on the frequency-adjustable oscillating Ramp type vortex generator is designed. Mach number M of incoming stream ∞ The internal shrinkage ratio of the inlet duct was 1.7, the length l=166 mm and the height h=19 mm of the isolation section=3.0. Oscillation height h of oscillation type Ramp type vortex generator capable of adjusting frequency v =4mm, boundary layer thickness δ=2h v Vortex generator thickness 0.5mm, spacing between centerlines s=7.5h=8 mm v The spreading position of the arranged frequency-adjustable oscillating Ramp type vortex generators is not more than the edge of the lip cover (only three vortex generators are arranged along the spreading direction due to the limitation of the width of a channel), the basic configuration of the oscillating Ramp type vortex generators is consistent with that of a classical Anderson type slope type vortex generator, and the slope chord length c=7.2h v The vortex generator oscillates up and down about the oscillation axis at a frequency of 66hz (as shown in fig. 5) =28.8mm. The distance between the trailing edge of the frequency-adjustable oscillating Ramp type vortex generator and the non-sticky incident position of the incident shock wave is 16 delta.
Parameter analysis is performed on the flow field of the vortex generator applied to the interference of shock waves/boundary layers of an inlet lip shroud, and as shown in table 1, the vibration type Ramp vortex generator and the fixed geometry vortex generator are x=10h away from the tail edge of the vortex generator v Throat boundary layer characteristic parameters at the point. It can be seen that the oscillatory Ramp type vortex generator has a reduced displacement thickness of 61.37% and a reduced momentum loss thickness of 18.46% and a reduced shape factor of 52.82% compared to the fixed geometry vortex generator. The vortex generator is obviously better in effect of inhibiting boundary layer separation than a fixed-geometry vortex generator. In FIG. 6, the boundary layer velocity profile changes at different moments in time at the cross-section of the oscillating Ramp type vortex generator at x=10hv from the trailing edge of the vortex generator are compared, and it can be seen that at any moment in time under the action of the oscillating Ramp type vortex generatorIs more filled than when t=0 (initial state).
The invention has many specific methods and approaches to implementing this solution, and the above description is only one embodiment of the invention. It should be noted that modifications and adaptations to the present invention may occur to one skilled in the art without departing from the principles of the present invention and are intended to be comprehended within the scope of the present invention. The components not explicitly described in this embodiment can be implemented by using the prior art.

Claims (8)

1. An air inlet channel design method based on a frequency-adjustable oscillating Ramp type vortex generator is characterized by comprising the following steps of:
(1) Placing an oscillation type Ramp type vortex generator capable of adjusting frequency on the surface of an inner channel of the air inlet channel, wherein the oscillation type Ramp type vortex generator is positioned at the upstream position of a non-sticky incidence point of a lip cover shock wave, and when the influence of separation on the performance of the air inlet channel is small, the vortex generator is embedded into the lower surface of the inner channel to keep the wall surface flat; when the separation has a relatively large influence on the performance of the air inlet channel, the vortex generator oscillates up and down through controllable frequency;
the air inlet based on the frequency-adjustable oscillating Ramp type vortex generator is a binary supersonic air inlet, and the working Mach number range is 0-3;
according to the oblique shock wave theory, mach number and pressure intensity of the lip cover after incidence and after reflection of the shock wave are calculated; the oblique shock theory is that the Mach number M 0 When supersonic airflow of the (a) encounters an object plane and deflects in the flowing direction delta, an oblique shock wave is emitted from the starting point of the turning of the object plane, the included angle between the wave plane and the airflow direction before the turning is an oblique shock wave angle beta, and the oblique shock wave angle beta and the Mach number M after the oblique shock wave can be obtained from the airflow deflection angle delta and the incoming flow Mach number 1 ；
Similarly, according to the intensity and the wave front Mach number of the lip cover incident laser wave, a criterion derived by adopting a free interference theory is expressed as a relation between pressure after separation and upstream parameters, and whether the inlet lip cover incident laser wave of the inlet channel under different incoming flow Mach numbers can cause obvious boundary layer separation is predicted;
wherein M is 0 Wave front Mach number, P, of incident shock wave for lip mask 0 For wave front static pressure of incident shock wave, P 1 For the static pressure value after reflection and excitation, C f，0 The wall friction coefficient is shown, and gamma is the adiabatic index of air;
(2) Frequency-adjustable oscillating Ramp type vortex generator size and mounting position selection criteria:
the arrangement direction of the vortex generator arranged on the lower surface of the inner channel of the air inlet channel is perpendicular to the supersonic incoming flow direction.
2. The design method according to claim 1, wherein the spacing between centerlines of the frequency-tunable oscillating Ramp type vortex generator is s=7.5 h v Wherein h is v Is the maximum height of the oscillation of the vortex generator.
3. The design method according to claim 1, wherein the basic configuration of the oscillating Ramp type vortex generator is basically consistent with that of a classical Anderson type slope type vortex generator, the vortex generator is isosceles triangle, the half apex angle AP of the vortex generator is 24 °, the thickness d of the vortex generator is 0.5mm, and the slope chord length c=7.2 h v The vortex generator swings up and down around the oscillation axis.
4. The design method of claim 1, wherein the development of increased momentum flux downstream of the tunable oscillating Ramp vortex generator is measured by: calculating a correlation coefficient R between the size of the separated bubbles and the momentum flux E to obtain an accurate R value, wherein the correlation coefficient R has an obvious maximum value at the distance of about Y=0.43 delta from the wall surface; it can be derived from the following equation that at a height y=0.43 δ or so, the vortex generator requires a fixed length of x=5.7 δ to allow the boundary layer to mix well to inject maximum momentum in the near wall region; delta is the boundary layer thickness;
wherein x is the distance from the downstream to the tail edge of the vortex generator, E (x) is the magnitude of momentum flux at different flow direction positions downstream of the vortex generator, U is the speed of the vortex generator, U clean For speed without vortex generator, U ∞ Is the far forward incoming flow velocity.
5. The method of claim 4, wherein the relationship between E (x) and x in the formula is as follows: x/δ=5.7 is preceded by a boundary layer intimate mixing region, the high momentum fluid is transported to the surface by the flow direction vortex, and the momentum flux E tends to increase; x/δ=5.7 followed by a plateau region of length 3.8δ, the momentum flux E becomes approximately constant; and the shock impact region evident at x/δ=9.5.
6. The design method according to claim 4, wherein a distance between a trailing edge point of the tunable oscillating Ramp vortex generator and a non-sticking incident point of an incident shock wave is greater than L; where l=7.8δ+5.7δ.
7. The design method according to claim 1, wherein the oscillation height of the frequency-adjustable oscillation type Ramp vortex generator is 10% -70% of the thickness of the local boundary layer.
8. The design method according to claim 1, wherein the oscillation height and frequency of the frequency-adjustable oscillating Ramp type vortex generator are adjusted according to the actual control requirement of the air inlet channel; comprises adjusting the frequency and the oscillation height of the up-and-down reciprocating oscillation of the vortex generator.
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Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101172520A (en) * 2006-10-31 2008-05-07 通用电气公司 Auxiliary power unit assembly
CN203114426U (en) * 2011-07-09 2013-08-07 拉姆金动力系统有限责任公司 Gas turbine engine and device for gas turbine engine
CN104863716A (en) * 2015-04-24 2015-08-26 南京航空航天大学 Design method for control measure of oblique shock wave/boundary layer interaction in air inlet on basis of binary bulge
CN107091157A (en) * 2017-06-05 2017-08-25 南京航空航天大学 A kind of imperial palace shrinkage ratio, determine geometry binary hypersonic inlet and design method
CN111561405A (en) * 2020-04-28 2020-08-21 中国航发湖南动力机械研究所 Double-duct pulse detonation engine and aircraft with same

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20170137116A1 (en) * 2009-07-10 2017-05-18 Peter Ireland Efficiency improvements for flow control body and system shocks
US9429071B2 (en) * 2011-06-23 2016-08-30 Continuum Dynamics, Inc. Supersonic engine inlet diffuser with deployable vortex generators

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101172520A (en) * 2006-10-31 2008-05-07 通用电气公司 Auxiliary power unit assembly
CN203114426U (en) * 2011-07-09 2013-08-07 拉姆金动力系统有限责任公司 Gas turbine engine and device for gas turbine engine
CN104863716A (en) * 2015-04-24 2015-08-26 南京航空航天大学 Design method for control measure of oblique shock wave/boundary layer interaction in air inlet on basis of binary bulge
CN107091157A (en) * 2017-06-05 2017-08-25 南京航空航天大学 A kind of imperial palace shrinkage ratio, determine geometry binary hypersonic inlet and design method
CN111561405A (en) * 2020-04-28 2020-08-21 中国航发湖南动力机械研究所 Double-duct pulse detonation engine and aircraft with same

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