Vortex compression regulated supersonic inlet

Abstract

The application belongs to the field of aviation technology and aircraft air intake and exhaust system, and particularly relates to a vortex compression adjusting supersonic air inlet, an outer cover and a bulge; the bulge is located on the surface of a machine body, the outer cover covers the bulge, and the bulge and the outer cover form an inlet of the air inlet; the outer shape surface parameters of the bulge are set so that the surface of the bulge forms a spanwise vortex flowing to both sides of the bulge, the actual compression angle of the bulge is smaller than the required theoretical compression angle under a set Mach number, and the difference between the actual compression angle and the theoretical compression angle is determined by the volume of the spanwise vortex, so that higher flight speed can be achieved even for the air inlet with a smaller fixed geometric compression angle.

Description

TECHNICAL FIELD

[0001] The application belongs to the field of aviation technology and aircraft air intake and exhaust system, and particularly relates to a vortex compression adjustable supersonic inlet. BACKGROUND

[0002] Supersonic aircraft powered by turbofan engine generally adopts an external pressure type supersonic inlet. By optimizing the compression angle of the inlet, a series of oblique shock waves and a terminal normal shock wave are generated on the compression surface to efficiently reduce the supersonic flow at the inlet entrance to subsonic speed and then provide the engine with the air in the inner duct. Generally, the maximum flow area of the inlet throat is required when the aircraft accelerates at subsonic speed and the engine is at maximum speed, while the maximum inlet entrance area and the minimum throat area are required when the aircraft flies at supersonic speed close to the maximum Mach number. When designing the inlet, the area of the throat must be allowed to pass the maximum flow of air required by the engine, and cannot be blocked, otherwise the engine thrust will be affected. The higher the flight speed of the aircraft, the greater the total compression angle of the compression surface is required, so that the oblique shock wave generated by the compression angle can reduce the Mach number in front of the normal shock wave to an acceptable range, reduce the loss of the normal shock wave, and avoid the separation of the boundary layer air flow caused by the excessive adverse pressure gradient of the normal shock wave.

[0003] According to the flow characteristics of the turbofan engine, the excess throat space can be transferred to the compression surface space required for increasing the compression angle when flying at supersonic speed, thereby generating a supersonic inlet with adjustable compression angle, which can minimize the unnecessary inlet entrance area and thus reduce the spillage resistance. However, the complex adjustment mechanism occupies valuable space of the aircraft, increases the weight of the aircraft, and increases the manufacturing and maintenance costs. Therefore, it is generally used in 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 and fixed-geometry supersonic external pressure inlet is generally used. Although the fixed-geometry inlet has the disadvantage of large entrance area and large resistance, it eliminates the cumbersome adjustment mechanism, and thus is more advantageous than the adjustment mechanism for light aircraft. Therefore, the fixed-geometry inlet is widely used in light aircraft.

[0004] There are two special fixed-geometry inlets, Caret and Bump inlets Figure 1 ), whose outer covers and compression surface leading edges are backward or forward swept. This special configuration limits the mechanical adjustment, and the longer inlet compression surface of the Caret and Bump inlets results in a larger inlet area and greater spillage resistance than the conventional non-stealth fixed-geometry supersonic inlet.

[0005] For the Bump inlet with the maximum flight Mach number less than 1.6, it is usually designed based on the attached flow assumption, and the partial fuselage boundary layer is removed out of the inlet by the compression bump spanwise pressure gradient, so that too much low-energy flow is avoided from entering the inlet to reduce the performance. Based on the flow characteristics of such attached flow, the bump can not only provide compression to reduce the Mach number in front of the normal shock and reduce the normal shock loss, but also can change the shock shape by the special shape of the bump, adjust the supersonic spill flow, etc.

[0006] With the increase of the free stream Mach number, the Mach number in front of the normal shock increases, the shock strength increases, and the pressure ratio behind and in front of the wave also increases, that is, the adverse pressure gradient in front of and behind the normal shock increases, so when the free stream Mach number exceeds a certain value, this adverse pressure gradient will cause the low-energy boundary layer flow to separate significantly, and it is no longer in the attached flow state. The degree of flow separation will increase with the increase of the Mach number until it affects the stable operation of the inlet, thereby limiting the increase of the maximum speed. SUMMARY

[0007] To solve the above problems, the application provides a vortex compression and supersonic inlet, comprising:

[0008] an outer cover and a bump; the bump is located on the surface of the fuselage, and the outer cover covers the bump, and the bump and the outer cover form the inlet of the inlet;

[0009] The outer shape parameters of the bump are set so that the surface of the bump forms a spanwise vortex flowing to both sides of the bump, the actual compression angle of the bump is smaller than the required theoretical compression angle at a set Mach number, and the difference between the actual compression angle and the theoretical compression angle is determined by the volume of the spanwise vortex.

[0010] Preferably, it further comprises a jet structure, which injects a jet into the spanwise vortex when the aircraft exceeds a set supersonic cruise Mach number, so as to avoid the spanwise vortex from spreading upstream to the conical shock and breaking down.

[0011] Preferably, the outer cover is swept forward, and the bump comprises a conical guide bump.

[0012] Preferably, the inlet belongs to a Bump inlet.

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

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

[0015] Preferably, the jet structure comprises jet slots distributed along the spanwise direction, the jet slots have an angle less than 45 degrees with the natural airflow direction, and the outlet of the jet slots is tangent to the airflow surface.

[0016] Preferably, the jet slots have valve pieces for controlling the flow size and switching.

[0017] The advantages of the present application include: compared with the flexible skin adjustable technology which has not yet been applied, the present application adopts pneumatic adjustment, saves the skin, is simple and effective, the vortex compression improves the total pressure recovery performance of the inlet and increases the stable working margin of the inlet, and can be directly applied to the inlet improvement of the active stealth aircraft and the supersonic inlet scheme design of the future aircraft. BRIEF DESCRIPTION OF DRAWINGS

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

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

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

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

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

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

[0024] In order to make the technical solutions of the present application and its advantages clearer, the technical solutions of the present application will be further described in detail below in combination with the drawings. It should be understood that the specific embodiments described herein are only part of the embodiments of the present application, and are only used to explain the present application, but not to limit the present application. It should be noted that, for the purpose of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design. In the case of no conflict, the embodiments in the present application and the technical features in the embodiments can be combined 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 surface will form a separation bubble under the counter-pressure action of the incident shock wave, as shown in Figure 2The two-dimensional separation bubble is shown in the schematic diagram, which is characterized by the entire flow field being divided into two fluid non-interacting areas: the separation zone upstream area and the backflow area. The flow lines in the separation zone upstream area cannot enter the backflow area, and the fluid in the backflow area does not come from the upstream flow area, but from a certain distance downstream. This separation is called closed separation. The outer boundary of the backflow area forms a new object surface, which lifts the boundary layer of the flow and continues to flow along the object surface. Due to the compression effect of the separation bubble on 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 with a certain design Mach number, as the incoming Mach number increases, the Mach number before the normal shock increases, and the shock strength continues to increase. When the incoming Mach number exceeds a certain value, the inverse pressure gradient between the wave front and the wave back will cause the boundary layer gas flow 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 of the separation flow volume increasing with the increase of the Mach number as self-adaptive adjustment with the Mach number. If this separation flow enters the inlet, it will have an adverse effect on the engine thrust. By adjusting the relevant parameters of the Bump inlet, the separation flow can be pushed to the two sides of the inlet and discharged from the inlet by means of the spanwise pressure difference of the inlet, thereby avoiding the adverse effects of the separation flow. According to the shock / boundary layer interaction theory mentioned above, this separation bubble belongs to closed separation, which is equivalent to lifting the object surface, thereby inducing a shock wave to provide compression. Since the flow behind the normal shock is subsonic, the adverse effects of the expansion wave and the reattachment shock wave can be eliminated.

[0027] The two-dimensional separation bubble is easily broken into the inlet under slightly larger back pressure, causing the inlet to work unstably, so the concept of three-dimensional vortex flow forming spanwise flow is proposed. Through the existing technology of the design of the bump or the bump and the outer cover together, three-dimensional vortex flow, i.e. spanwise vortex flow 2, can be generated, and the outer shape surface of the bump 6 is set so that the surface of the bump 6 forms spanwise vortex flow 2 to the two sides of the bump 6, as shown in Figure 3 and Figure 4 .

[0028] The three-dimensional separation vortex flow and the two-dimensional separation bubble have very similar shapes in the two-dimensional cross section, but the flow characteristics are very different. The three-dimensional separation vortex flow has the backflow motion characteristics of the two-dimensional space of the separation bubble, and also has flow velocity along the spanwise direction (vortex axis direction). In three-dimensional space, air particles exhibit a spiral motion pattern, as shown in Figure 4 , the spanwise flow is a basic condition for maintaining the stability of the three-dimensional separation vortex flow, and is also a basic condition for the total pressure recovery performance of the inlet to benefit.

[0029] The large inlet area of the inlet duct will cause large spillage drag, which will limit the maximum use Mach number of the inlet duct. For the bulge 6 inlet duct of the forward swept cowl 7, the bulge 6 is of the conical shape. In the case of the same design Mach number, the inlet area of the inlet duct is reduced by reducing the compression angle of the bulge 6, so as to reduce the spillage drag. The higher counter pressure behind the normal shock wave outside the inlet duct inlet is used to separate the boundary layer flow on the surface of the bulge 6, so as to form the spanwise vortex flow 2 flowing to both sides of the bulge 6. The low-energy flow of the boundary layer is completely discharged, so that the inlet duct has higher total pressure recovery performance. At the same time, the space volume occupied by the vortex flow to both sides also provides additional compression capacity, so as to reduce the requirement of the fixed geometric compression angle.

[0030] The volume of the three-dimensional separated vortex flow is related to the strength of the normal shock wave, and the strength of the normal shock wave is proportional to the Mach number of the wave front. This means that when the Mach number of the incoming flow increases, the normal shock wave is enhanced, the counter pressure behind the wave is increased, the volume of the separated vortex flow is increased, and the auxiliary compression capacity provided is improved, as shown by the curve. Figure 5 As shown by the curve, the total pressure recovery coefficient is increased with the increase of the counter pressure, which is the contribution of the three-dimensional separated vortex flow. The auxiliary compression which is self-adapted to the Mach number of the incoming flow is similar to the regulation of the conventional adjustable inlet duct. The stability of the three-dimensional separated vortex flow is determined by the spanwise outflow capacity of the vortex flow. When the volume of the three-dimensional separated vortex flow spreads upstream to the vicinity of the conical shock wave 5, the instability phenomenon of vortex flow breakage will occur, and part of the low-energy flow will begin to enter the pipeline, so that the total pressure recovery at the outlet of the inlet duct begins to decrease, and further into the surge state. Therefore, additional flow control measures are needed to prevent the vortex flow from spreading upstream. One of the methods is to open a slot along the span of the bulge 6, and the outlet of the slot is tangent to the airflow surface, as shown in Figure 3 By the jet flow in the direction of the airflow, the degree of upstream spread of the separated vortex flow is slowed down, so as to avoid the occurrence of surge. When the aircraft flies below the supersonic cruise Mach number, the inlet duct can be set not to apply the jet flow; when the aircraft exceeds the supersonic cruise Mach number, the jet flow can be turned on, so that the inlet duct can adapt to a higher flight Mach number under the vortex auxiliary compression.

[0031] In summary, the spanwise vortex flow 2 is the basic flow configuration of the inlet duct design, and the jet flow slot 1 is designed at a specific position. By stabilizing the large volume of separated vortex flow through the jet flow, the inlet duct with a smaller fixed geometric compression angle can adapt to a higher flight speed.

[0032] The above is only a specific embodiment of the present application, but the protection scope of the present application is not limited thereto. Any changes or replacements within the technical range disclosed in the present application can be easily thought by those skilled in the art, which should be covered within the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A supersonic inlet with vortex compression regulation, characterized in that, include: The outer cover and the 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 the inlet of the air intake; The surface parameters of the bulge are set such that spanwise vortices are formed on the surface of the bulge and flow toward both sides of the bulge. The actual compression angle of the bulge is less than the theoretical compression angle required at 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-regulated supersonic air intake as described in claim 1, characterized in that, It also includes a jet structure that injects a jet into the spanwise vortex when the aircraft exceeds the set supersonic cruise Mach number, in order to prevent the spanwise vortex from breaking up and becoming unstable when its volume spreads upstream to the conical shock wave.

3. The vortex compression regulating supersonic air intake as described in claim 1, characterized in that, The outer cover sweeps forward, and the bulge includes a cone-shaped guide bulge.

4. The vortex compression regulating supersonic air intake as described in claim 1, characterized in that, The air intake is a bump intake.

5. The vortex compression regulating supersonic air intake as described in claim 1, characterized in that, The spanwise vortex has a spanwise vortex axis, and the air particles in the spanwise vortex move spirally around the vortex axis from the center of the bulge to both sides of the spanwise direction.

6. The vortex compression-regulated supersonic air intake as described in claim 1, characterized in that, The spanwise vortex is generated by the interference between the normal shock wave of the air intake and the boundary layer of the air intake.

7. The vortex compression regulating supersonic air intake as described in claim 2, characterized in that, The jet structure includes jet slots distributed along the spanwise direction, the jet slots being located at the location of the spanwise vortex, and the angle between the airflow direction of the jet slots and the natural incoming flow direction being less than 45 degrees.

8. The vortex compression regulating supersonic air intake as described in claim 7, characterized in that, The jet slit has a valve plate that controls the flow rate and allows for on / off switching.

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