Vortex compression adjusting supersonic air inlet channel

Abstract

The invention belongs to the field of aviation technologies and airplane air inlet and exhaust systems, and particularly relates to a vortex compression adjusting supersonic air inlet channel which comprises an outer cover and a bulge. The bulge is located on the surface of the machine body, the outer cover covers the bulge, and an inlet of the air inlet channel is formed by the bulge and the outer cover; the parameters of the outer profile of the bump are set to enable the surface of the bump to form spanwise vortexes flowing to the two sides of the bump, the actual compression angle of the bump is smaller than a theoretical compression angle required under a set Mach number, and the difference value between the actual compression angle and the theoretical compression angle is determined by the volume of the spanwise vortexes; and a higher flight speed can be realized even under an air inlet channel with a smaller fixed geometric compression angle.

Description

Technical Field

[0001] The present application belongs to the field of aviation technology and aircraft intake and exhaust systems, and in particular relates to a vortex compression regulated supersonic inlet. Background Art

[0002] Supersonic aircraft powered by aviation turbofan engines generally use external pressure supersonic inlets. By optimizing the compression angle of the inlet, the compression surface generates a series of oblique shock waves and terminal normal shock waves to efficiently reduce the supersonic incoming flow to subsonic speed outside the inlet inlet, and then enter the inner pipe to provide it to the engine. Usually, when the aircraft accelerates at subsonic speed and the engine is at the maximum speed, the required inlet throat flow area is the largest, while when flying at supersonic speed close to the maximum Mach number, the required inlet inlet area is the largest and the throat area is the smallest. When designing the inlet, it is first necessary to ensure that the throat area allows the maximum flow of air required by the engine to pass through, and it cannot be blocked, otherwise it will affect the engine thrust. The higher the flight speed of the aircraft, the greater the total compression angle of the inlet compression surface required, so that the oblique shock wave generated by the compression angle can reduce the Mach number before the normal shock wave to an acceptable range, reduce the loss of the normal shock wave, and avoid the boundary layer airflow separation caused by the excessive reverse pressure gradient of the normal shock wave.

[0003] In view of the flow characteristics of aviation turbofan engines, during supersonic flight, the excess throat space can be transferred to the compression surface space required to increase the compression angle, thus giving birth to a supersonic inlet with adjustable compression angle, which can minimize the unnecessary inlet inlet area, thereby reducing overflow resistance. However, the complex adjustment mechanism occupies the valuable space of the aircraft, increases the aircraft weight, manufacturing and maintenance costs, and is generally used on heavy aircraft with a maximum flight Mach number greater than 2.0. For aircraft with a maximum flight Mach number less than 2.0, a relatively simple, geometrically fixed supersonic external pressure inlet is usually used. Although the fixed geometry inlet brings the disadvantages of a large inlet area and high resistance, the removal of the bulky adjustment mechanism has more advantages than disadvantages for light aircraft, and therefore, it is widely used by light aircraft.

[0004] There are two special fixed geometry inlet ports: Caret and Bump. Figure 1 ), their outer cover and compression surface front edges are swept back or forward. This special configuration determines that its mechanical adjustability is limited. At the same time, its longer inlet compression surface causes the inlet area to be larger than the conventional non-stealth fixed geometry supersonic inlet inlet area, and the overflow resistance is greater.

[0005] For bump inlets with a maximum flight Mach number less than 1.6, they are usually designed based on the assumption of attached flow, using the compression of the spanwise pressure gradient of the bump to remove part of the fuselage boundary layer from the inlet, avoiding excessive low-energy flow entering the inlet and reducing performance. Based on the flow characteristics of such attached flow, in addition to providing compression to reduce the Mach number before the normal shock wave and reduce the normal shock wave loss, the special shape of the bump can also be used to change the shock wave shape, adjust the supersonic overflow, and so on.

[0006] For this geometrically fixed inlet, as the incoming flow Mach number increases, the Mach number of the positive shock wave front increases, the shock wave intensity also increases, and the pressure ratio between the wave rear and the wave front also increases, that is, the adverse pressure gradient before and after the positive shock wave increases. Therefore, when the free incoming flow Mach number exceeds a certain value, this adverse pressure gradient will cause the low-energy boundary layer airflow to separate significantly, and it will no longer be an attached flow state. The degree of flow separation will increase with the increase of Mach number until it affects the stable operation of the inlet, thereby limiting the increase in maximum speed. Summary of the invention

[0007] In order to solve the above problems, the present application provides a vortex compression regulating supersonic inlet, comprising:

[0008] An outer cover and a bulge; the bulge is located on the surface of the body, the outer cover covers the bulge, and the bulge and the outer cover form an inlet of the air intake;

[0009] The parameters of the bulge outer surface are set so that spanwise vortices are formed on the bulge surface flowing to both sides of the bulge, and the actual compression angle of the bulge is smaller than the theoretical compression angle required under the set Mach number. The difference between the actual compression angle and the theoretical compression angle is determined by the volume of the spanwise vortex.

[0010] Preferably, it also includes a jet structure, which injects a jet into the spanwise vortex when the aircraft exceeds a set supersonic cruise Mach number to prevent the spanwise vortex from breaking and becoming unstable when the volume of the spanwise vortex spreads upstream to the conical shock wave.

[0011] Preferably, the outer cover is forward swept and the bulge comprises a cone-guided bulge.

[0012] Preferably, the air intake duct is a bump air intake duct.

[0013] Preferably, the spanwise vortex has a spanwise vortex axis, and air particles of the spanwise vortex move from the middle of the bulge to both sides of the spanwise bulge in a spiral manner around the vortex axis.

[0014] Preferably, the spanwise vortex is generated by interference between a normal shock wave of the inlet and a boundary layer of the inlet.

[0015] Preferably, the jet structure comprises jet slits distributed along the span direction, the angle between the airflow direction of the jet slits and the natural flow direction is less than 45 degrees, and the outlet of the jet slit is tangent to the airflow surface.

[0016] Preferably, the jet slit has a valve plate for controlling the flow rate and switching.

[0017] The advantages of the present application include: compared with the flexible skin adjustable technology that has not yet been applied, the present invention adopts aerodynamic adjustment, eliminates the skin, is simple and effective, and the vortex compression improves the total pressure recovery performance of the inlet while also increasing the stable working margin of the inlet. It can be directly applied to the improvement of the inlet of existing stealth aircraft and the design of supersonic inlet schemes for future aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 This is a schematic diagram of the Bump air inlet and jet control in the preferred embodiment of the present application;

[0019] Figure 2 is a schematic diagram of the separation bubble produced by shock wave / boundary layer interference;

[0020] Figure 3 It is a schematic diagram of the spanwise vortex cross section of the present application;

[0021] Figure 4 It is a three-dimensional schematic diagram of spanwise flow vortex of the present application;

[0022] Figure 5 This is a schematic diagram of the relationship between the spanwise flow vortex volume and performance of this application.

[0023] 1-jet gap, 2-spanwise vortex, 3-normal shock wave, 4-shock wave generated by spanwise vortex, 5-conical shock wave, 6-bulge, 7-outer cover. DETAILED DESCRIPTION

[0024] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be further described in detail in detail and in detail with reference to the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0025] According to the shock wave / boundary layer interference theory, the low-energy flow in the boundary layer close to the object surface will form a separation bubble under the back pressure of the incident shock wave, such as Figure 2The schematic diagram of the two-dimensional separation bubble shown in the figure is characterized by the fact that the entire flow field is divided into two areas where the fluids do not enter each other: the upstream area of ​​the separation zone and the recirculation zone. The streamlines in the upstream area of ​​the separation zone cannot enter the recirculation zone, and the fluid in the recirculation zone does not come from the upstream flow zone, but flows in from a certain distance downstream. This separation is called closed separation. The outer boundary of the recirculation zone is equivalent to forming a new surface, and the boundary layer of the future flow is lifted and bypassed and continues to flow on the surface. Because the bulge of the separation bubble compresses the supersonic incoming flow, an induced separation shock wave, an expansion wave and a reattachment shock wave are formed.

[0026] For a fixed geometry bump inlet designed with a certain Mach number, as the incoming flow Mach number increases, the Mach number of the positive shock wave increases, and the shock wave intensity continues to increase. When the incoming flow Mach number exceeds a certain value, the adverse pressure gradient between the wave front and the wave rear will cause the boundary layer airflow that has not been completely drained on the bump 6 to separate, and the volume of this separation will increase with the increase of the Mach number. We call this change in the volume of the separated flow increasing with the increase of the Mach number as adaptive regulation with the Mach number. If this separated flow enters the inlet, it will have an adverse effect on the engine thrust. By adjusting the relevant parameters of the bump inlet, the separated flow can be pushed to both sides of the inlet and discharged from the inlet with the help of the inlet spanwise pressure difference, thereby avoiding the adverse effects of the separated flow. According to the shock wave / boundary layer interference theory mentioned above, this separation bubble belongs to a closed separation, which is equivalent to raising the object surface, thereby inducing the generation of shock waves and providing compression. Since the positive shock wave is followed by a subsonic airflow, the adverse effects of the expansion wave and the reattached shock wave can be eliminated.

[0027] The two-dimensional separation bubble is easy to break and enter the air inlet under a slightly larger back pressure, causing the air inlet to work unstably. Therefore, the concept of forming a three-dimensional vortex of spanwise flow is proposed. Through the prior art design of the bulge or the bulge and the outer cover, a three-dimensional vortex, i.e., a spanwise vortex 2, can be generated. The outer surface of the bulge 6 is configured so that the surface of the bulge 6 forms a spanwise vortex 2 flowing to both sides of the bulge 6, such as Figure 3 and Figure 4 shown.

[0028] The three-dimensional separation vortex and the two-dimensional separation bubble are very similar in appearance in the two-dimensional cross section, but the flow characteristics are very different. The three-dimensional separation vortex has the backflow motion characteristics of the separation bubble in the two-dimensional space, and also has flow velocity along the span direction (vortex axis direction). In the three-dimensional space, the air particles show a spiral motion pattern. The macroscopic flow appearance is as follows: Figure 4 As shown in the figure, spanwise flow is the basic condition for maintaining the stability of the three-dimensional separated vortex and also the basic condition for benefiting the inlet total pressure recovery performance.

[0029] A large inlet inlet area will cause large overflow resistance, thereby limiting the maximum usable Mach number of the inlet. For the bulge 6 inlet of the forward-swept outer cover 7, the bulge 6 is in a cone-guide shape. Under the condition that the design Mach number remains unchanged, the inlet area of ​​the inlet is reduced by reducing the compression angle of the bulge 6, thereby reducing the overflow resistance, and using the higher back pressure after the positive shock wave is pushed out of the inlet inlet to separate the boundary layer airflow on the surface of the bulge 6, forming a spanwise vortex 2 flowing to both sides of the bulge 6, which completely discharges the low-energy airflow in the boundary layer and enables the inlet to obtain a higher total pressure recovery performance. At the same time, the space volume occupied by the vortex flowing to both sides also provides additional compression capacity, thereby reducing the need for a fixed geometric compression angle.

[0030] The volume of the three-dimensional separation vortex is related to the strength of the normal shock wave, and the strength of the normal shock wave is proportional to the wavefront Mach number. This means that when the incoming flow Mach number increases, the normal shock wave is strengthened, the back pressure behind the wave increases, the volume of the separation vortex increases, and the auxiliary compression capacity provided is improved, such as Figure 5 As shown in the curve, the total pressure recovery coefficient increases with the increase of back pressure, which is the contribution of the three-dimensional separated vortex. This auxiliary compression that is adaptively adjusted to the incoming flow Mach number is similar to the regulation law of the conventional adjustable inlet. The stability of the three-dimensional separated vortex is determined by the spanwise outflow capability of the vortex. When the volume of the three-dimensional separated vortex spreads upstream to approach the conical shock wave 5, the vortex breakup instability phenomenon will occur, and some low-energy flow will begin to enter the pipeline, causing the total pressure recovery at the inlet outlet to begin to decline, and further enter a surge state. Therefore, additional flow control measures are needed to prevent the vortex from spreading upstream. One method is to make a slit along the span of the bulge 6, and the outlet of the slit is tangent to the airflow surface, such as Figure 3 As shown in the figure, the jet in the direction of the airflow slows down the spread of the separation vortex upstream, thereby avoiding the occurrence of surge. When the aircraft flies below the supersonic cruise Mach number, the inlet can be set not to apply the jet; when the aircraft exceeds the supersonic cruise Mach number, the jet can be turned on so that the inlet can adapt to a higher flight Mach number under vortex-assisted compression.

[0031] In summary, the spanwise vortex 2 is used as the basic flow configuration for the inlet design, and the jet slot 1 is designed at a specific position. By stabilizing the separation vortex of a larger volume through the jet, the inlet with a smaller fixed geometric compression angle can be adapted to a higher flight speed.

[0032] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A vortex compression regulated supersonic inlet, characterized in that: include: An outer cover and a bulge; the bulge is located on the surface of the body, the outer cover covers the bulge, and the bulge and the outer cover form an inlet of the air intake; The parameters of the bulge outer surface are set so that spanwise vortices are formed on the bulge surface flowing to both sides of the bulge, and the actual compression angle of the bulge is smaller than the theoretical compression angle required under the set Mach number. The difference between the actual compression angle and the theoretical compression angle is determined by the volume of the spanwise vortex.

2. The vortex compression regulating supersonic inlet according to claim 1, characterized in that: It also includes a jet structure, which injects a jet into the spanwise vortex when the aircraft exceeds a set supersonic cruise Mach number to prevent the spanwise vortex from breaking and becoming unstable when the volume of the spanwise vortex spreads upstream to the conical shock wave.

3. The vortex compression regulating supersonic inlet according to claim 1, characterized in that: The outer cover is swept forward, and the bulge includes a cone-guided bulge.

4. The vortex compression regulating supersonic inlet according to claim 1, characterized in that: The air intake duct is a bump air intake duct.

5. The vortex compression regulating supersonic inlet according to claim 1, characterized in that: The spanwise vortex has a spanwise vortex axis, and air particles of the spanwise vortex move from the middle of the bulge to both sides of the bulge in a spiral manner around the vortex axis.

6. The vortex compression regulating supersonic inlet according to claim 1, characterized in that: The spanwise vortex is generated by the interference between the normal shock wave of the inlet and the boundary layer of the inlet.

7. The vortex compression regulating supersonic inlet according to claim 2, characterized in that: The jet structure comprises jet slots distributed along the span direction, the jet slots are located at the positions of the span direction vortices, and the angle between the airflow direction of the jet slots and the natural incoming flow direction is less than 45 degrees.

8. The vortex compression regulating supersonic inlet according to claim 2, characterized in that: The jet slit is provided with a valve sheet for controlling the flow rate and switching.

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