Active jet control method for forced transition of boundary layer in ultra-high speed inlet

Abstract

The present invention provides an active jet control method for forced transition of boundary layer in an ultra-high-speed inlet, wherein a row of equally spaced nozzles are arranged on a front body or an inlet compression surface of an air-breathing ultra-high-speed aircraft in a direction perpendicular to the flow, a gas source is ejected from the nozzles through a delivery pipeline, the ejected gas interacts with the incoming air to generate a flow vortex structure, and the ultra-high-speed boundary layer is induced to be forced to transition to turbulence; the control method comprises determining the installation position of the nozzles, determining the injection medium, determining the nozzle diameter, determining the total pressure before injection, determining the nozzle spacing and the number of nozzles; compared with the existing passive forced transition control method, the method can significantly widen the control range of forced transition of the boundary layer, reduce the loss of inlet performance caused by control measures, and avoid direct collision between the transition zone and the high-speed airflow, and basically does not require additional thermal protection design, which is of great significance for ensuring the wide range and stable operation of the air-breathing ultra-high-speed aircraft.

Description

Technical Field

[0001] The present invention relates to the field of air-breathing hypersonic aircraft and engines, and particularly to the design of air inlet flow control. Background Art

[0002] For air-breathing hypersonic aircraft, if the fluid in the boundary layer of the inlet is laminar, it is easy to cause the flow separation in the compression corner and the lip shock wave reflection area. This separation has great uncertainty and may cause the inlet to fail to start and the aircraft to lose control, thus posing a serious safety hazard to the aircraft; but if the fluid entering the inlet turns into turbulence, the boundary layer's ability to resist adverse pressure gradients can be greatly enhanced, thereby effectively suppressing this separation and facilitating the start of the inlet. In fact, when a small-sized air-breathing hypersonic aircraft flies at high altitudes, due to the low turbulence and noise of the incoming flow (less than 0.1%), the boundary layer in the inlet is often difficult to naturally transition, resulting in safety hazards in flight. Therefore, it is necessary to install a control device on the forebody to promote the forced transition of the boundary layer to turbulence, increase the boundary layer's ability to resist adverse pressure gradients, reduce flow separation, and ensure the start of the inlet.

[0003] The existing boundary layer forced transition control mainly adopts passive forced transition devices, which can be divided into distributed transition zones and discrete transition zones. Distributed transition zones are mostly made of very small solid particles (such as corundum, etc.), and the size of the particles is very small and difficult to distinguish; while discrete transition zones are mostly made of metal particles, the particle size is generally about a few millimeters, and the particle shape is generally cylindrical, diamond-shaped, triangular, swept slope type, etc. These passive control devices have the advantages of simple structure, obvious control effect, stability and reliability, but their disadvantages are also very obvious: they can often only be used for a narrow range of flight envelopes. When the incoming flow conditions deviate from the design point, the forced transition effect will either fail due to the low height of the coarse particles, which will bring safety hazards to the aircraft, or the coarse particles will cause unnecessary loss of inlet performance due to the high height, and bring additional thermal protection requirements.

[0004] In order to overcome the shortcomings of the passive forced transition zone, the present invention provides a boundary layer forced transition control method of active jet, which can realize wide-range low-loss control of boundary layer forced transition in a wider range of flight Mach number and altitude. Summary of the invention

[0005] The purpose of the present invention is to provide an active jet control method for forced transition of the boundary layer of an ultra-high-speed inlet using wall jets. Compared with the existing passive forced transition control method, this method can significantly widen the control range of the boundary layer forced transition, reduce the performance loss of the inlet caused by control measures, and avoid the direct collision between the transition zone and the high-speed airflow. Basically, no additional thermal protection design is required, which is of great significance for ensuring the wide range and stable operation of air-breathing ultra-high-speed aircraft.

[0006] The above technical objectives of the present invention are achieved through the following technical solutions:

[0007] An active jet control method for forced transition of a boundary layer in a hyperspeed inlet is provided, wherein a row of equally spaced spray holes are arranged on the front body or the compression surface of the inlet of an air-breathing hyperspeed aircraft in a direction perpendicular to the flow direction, and a gas source is ejected from the spray holes through a delivery pipeline, and the ejected gas interacts with the incoming air to generate a flow vortex structure, thereby inducing forced transition of the hyperspeed boundary layer into turbulence;

[0008] The control method includes determining the installation position of the spray hole, determining the spray medium, determining the diameter of the spray hole, determining the total pressure before spraying, determining the spray hole spacing and the number of spray holes, and includes the following steps:

[0009] (1) Determine the nozzle installation position. The nozzle is installed on the front body or the compression surface of the air intake of the air-breathing hypersonic vehicle;

[0010] For multi-wave compression inlet, the nozzle is installed on the first compression surface.

[0011] For the three-dimensional curved compression inlet, the nozzle is installed on the front body to ensure that the streamwise vortex induced by the jet flow has enough instability distance;

[0012] At the same time, the nozzle installation position also needs to consider whether there is enough space in the aircraft to place the gas pipeline. After determining the installation position, the laminar boundary layer thickness δ at the installation position is obtained through fluid mechanics calculation in the laminar state. Here, the boundary layer thickness δ is defined as the total enthalpy boundary layer thickness. The total enthalpy at the outer edge of the boundary layer is 99% of the total enthalpy of the incoming flow.

[0013] (2) Determine the injection medium. The injection medium shall be air or nitrogen, and the gas storage pressure shall not be less than 1 atmosphere;

[0014] (3) Determine the nozzle diameter of 0.3-1mm and the total pressure before spraying P j The nozzle shape is circular and perpendicular to the compression surface, so that the jet airflow is perpendicular to the airflow in the boundary layer; the airflow penetration depth hIt is taken as 0.5 times the thickness of the local boundary layer. The airflow penetration depth is defined as: Get the flow field Mach number distribution on the center line of the vertical nozzle, and define the distance from the point on the center line of the vertical nozzle when the Mach number first reaches the local maximum to the wall as the penetration depth h , since the air flow penetration depth satisfies the relationship: ,in p 2 is the static pressure of the sonic jet airflow, p is the local fluid static pressure when there is no injection, and the parameter a , b Calibration is performed through fluid mechanics calculations to obtain parameters a , b Then, substitute into the relational equation to obtain the static pressure of the jet flow p 2. According to the isentropic relationship ,in γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ;

[0015] (4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is based on the penetration depth. h The number of installed nozzles is determined by the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The width w of the fluid entering the inlet duct is obtained through the wall limit streamlines. The width w at the nozzle installation position is the effective width range of the nozzle. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w / d.

[0016] As a preferred method, if there is already a nitrogen tank in the aircraft, no additional gas tank is needed. It is only necessary to introduce a pipeline from the nitrogen tank to connect it to the spray hole, and then control it through the solenoid valve and the pressure reducing valve.

[0017] As a preferred method, if there is no existing nitrogen gas source in the aircraft, air is used as the injection medium, and an additional gas tank is equipped, and the gas storage pressure of the gas tank is not less than 1 atmosphere.

[0018] As a preferred embodiment, the diameter of the nozzle hole is 0.5 mm.

[0019] By adopting the above technical solution, the gas enters the compression surface boundary layer from the nozzle at a certain injection pressure, interacts with the incoming flow, and generates a streamwise vortex, which induces the boundary layer to forcibly transform into turbulence.

[0020] Compared with the existing passive boundary layer forced transition control technology, the present invention has the following beneficial effects:

[0021] 1. Reduce the thermal protection requirements of the forced transition device. The present invention realizes forced transition by jetting airflow from the precursor or the wall of the air inlet, avoiding the particle structure protruding from the wall in the passive forced transition device, thus avoiding the high-speed collision process between the high-speed airflow and the particle structure, and no additional aerodynamic heating is generated, thus significantly reducing the thermal protection requirements of the transition device.

[0022] 2. Effectively broaden the working range. The effect of forced transition is seriously affected by the ratio of the height of coarse particles or the penetration depth to the thickness of the local boundary layer. The height of the existing passive forced transition device is fixed. When the incoming flow conditions deviate from the design point, if the thickness of the local boundary layer increases significantly, the forced transition device may fail. If the thickness of the local boundary layer decreases significantly, the coarse particles will contact the mainstream due to the high particle height, which will introduce large aerodynamic losses and cause large aerodynamic heating. The boundary layer forced transition active jet control method adopted by the present invention can achieve real-time adjustment of the penetration depth by setting a flow regulating valve in the air pipeline after the aircraft deviates from the design point. Combined with the fluid mechanics calculation before flight, the pre-spray pressure can be set according to the established timing, and the flight status can be judged by the sensor data during flight, and the pre-spray pressure can be adjusted in real time, thereby realizing a wide range of low-loss forced transition control.

[0023] 3. It provides a feasible boundary layer forced transition control scheme for wide-range air-breathing ultra-high-speed aircraft. The present invention can achieve wide-range low-loss control of boundary layer forced transition, can use nitrogen from the fuel supply system as the gas source, connect the gas storage tank and the spray hole through the gas pipeline, and control the flow through the solenoid valve, which is in line with my country's current industrial level and technical foundation and can be used in actual flight. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Figure 1 It is a schematic diagram of the multi-wave compression inlet of the present invention.

[0025] Figure 2 This is the process in which the jet gas interacts with the incoming flow to generate a streamwise vortex and gradually becomes unstable and breaks up.

[0026] Figure 3 It is a schematic diagram of the wall heat flux comparison and transition position on the compression surface of the inlet duct of the present invention.

[0027] In the figure, 1 is a nozzle. DETAILED DESCRIPTION

[0028] The following describes the embodiments of the present invention through specific examples, and those skilled in the art can easily understand other advantages and effects of the present invention from the contents disclosed in this specification. The present invention can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present invention.

[0029] Example 1

[0030] The embodiment provides an active jet control method for forced transition of a boundary layer in a hyperspeed inlet, wherein a row of equally spaced spray holes are arranged on a front body or a compression surface of an air-breathing hyperspeed aircraft in a direction perpendicular to the flow, and a gas source is ejected from the spray holes through a delivery pipeline, and the ejected gas interacts with the incoming air to generate a flow vortex structure, thereby inducing a forced transition of the hyperspeed boundary layer into turbulence;

[0031] The control method includes determining the installation position of the spray hole, determining the spray medium, determining the diameter of the spray hole, determining the total pressure before spraying, determining the spray hole spacing and the number of spray holes, and includes the following steps:

[0032] (1) Determine the nozzle installation position. The nozzle is installed on the front body or the compression surface of the air intake of the air-breathing hypersonic vehicle;

[0033] For multi-wave compression inlet, the nozzle is installed on the first compression surface.

[0034] For the three-dimensional curved compression inlet, the nozzle is installed on the front body to ensure that the streamwise vortex induced by the jet flow has enough instability distance;

[0035] At the same time, the nozzle installation position also needs to consider whether there is enough space in the aircraft to place the gas pipeline. After determining the installation position, the laminar boundary layer thickness δ at the installation position is obtained through fluid mechanics calculation in the laminar state. Here, the boundary layer thickness δ is defined as the total enthalpy boundary layer thickness. The total enthalpy at the outer edge of the boundary layer is 99% of the total enthalpy of the incoming flow.

[0036] (2) Determine the injection medium. The injection medium shall be air or nitrogen, and the gas storage pressure shall not be less than 1 atmosphere;

[0037] (3) Determine the nozzle diameter of 0.3-1mm and the total pressure before spraying P j The nozzle shape is circular and perpendicular to the compression surface, so that the jet airflow is perpendicular to the airflow in the boundary layer; the airflow penetration depth h It is taken as 0.5 times the thickness of the local boundary layer. The airflow penetration depth is defined as: Get the flow field Mach number distribution on the center line of the vertical nozzle, and define the distance from the point on the center line of the vertical nozzle when the Mach number first reaches the local maximum to the wall as the penetration depth h, since the air flow penetration depth satisfies the relationship: ,in p 2 is the static pressure of the sonic jet airflow, p is the local fluid static pressure when there is no injection, and the parameter a , b Calibration is performed through fluid mechanics calculations to obtain parameters a , b Then, substitute into the relational equation to obtain the static pressure of the jet flow p 2. According to the isentropic relationship ,in γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ;

[0038] (4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is based on the penetration depth. h The number of installed nozzles is determined by the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The width w of the fluid entering the inlet duct is obtained through the wall limit streamlines. The width w at the nozzle installation position is the effective width range of the nozzle. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w / d.

[0039] In some embodiments, if there is already a nitrogen tank in the aircraft, no additional gas tank is needed. It is only necessary to introduce a pipeline from the nitrogen tank to connect to the spray hole, and then control it through the solenoid valve and the pressure reducing valve.

[0040] In some embodiments, if there is no existing nitrogen gas source in the aircraft, air is used as the injection medium, and an additional gas tank is provided, and the gas storage pressure of the gas tank is not less than 1 atmosphere.

[0041] In some embodiments, the nozzle hole diameter is 0.5 mm.

[0042] Example 2

[0043] In order to verify the effectiveness of the active jet control method for forced transition of ultra-high-speed boundary layer proposed in the present invention, a 4-wave compression inlet configuration (see Figure 1 ) designed a forced transition scheme and performed numerical simulations on it.

[0044] The inlet model is 550mm long, 225.7mm wide and 113mm high. The design Mach number range of the inlet is 4 to 7, the angle of attack is -1 to 3 degrees, and the yaw angle is less than 2 degrees. The starting Mach number is 4. The design state is: Mach number 6, angle of attack 1°, yaw angle 0°, total pressure of incoming flow 2Mpa, total temperature of incoming flow 455K, and unit Reynolds number of incoming flow is 18.9*10 6 / m, the incoming flow is pure air, and the steps for determining the boundary layer forced transition active jet control scheme are as follows:

[0045] (1) Determine the installation position of the nozzle. The inlet is a multi-wave compression inlet. The nozzle installation position is located on the first compression surface. Considering the internal pipeline layout requirements of the fuselage, the center of the nozzle is determined to be 88mm away from the leading edge of the inlet. Using laminar fluid mechanics calculations, the thickness of the laminar boundary layer at 88mm from the leading edge is δ=1.2mm, and the static pressure here is about 2850Pa.

[0046] (2) Determine the injection medium. This configuration is only an intake configuration, without a fuel supply system, and the injection medium is air.

[0047] (3) Determine the nozzle diameter of 0.3-1mm and the total pressure before spraying P j The nozzle shape is circular and perpendicular to the compression surface. The nozzle diameter is 0.5mm. Since the air flow penetration depth satisfies the relationship: ,in p 2 is the static pressure of the sonic jet airflow, p is the local fluid static pressure when there is no injection, and the parameter a , b Calibration is performed through fluid mechanics calculations to obtain parameters a , b Then, substitute into the relational equation to obtain the static pressure of the jet flow p 2. According to the isentropic relationship ,in γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ;

[0048] Computational fluid dynamics was used to calculate the working conditions under three different injection static pressures, and the penetration depth was calculated to obtain the formula , according to this formula, set =3.0, the corresponding penetration depth is 0.6mm, and the pre-spray pressure is .

[0049] (4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is based on the penetration depth. hThe number of installed nozzles is determined by the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The width w of the fluid entering the inlet duct is obtained through the wall limit streamlines. The width w at the nozzle installation position is the effective width range of the nozzle. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w / d.

[0050] The nozzle spacing is selected as 2.5mm, then d / h≈4.2, which is within the range required by the present invention. According to the wall limit streamline obtained by computational fluid dynamics, the nozzle action range is obtained to be within 36mm on both sides of the center line, so the total width of the action range is 72mm, and the nozzle spacing is 2.5mm, so a total of 29 nozzles are required.

[0051] Numerical simulation was used to obtain the flow field structure, transition position and wall heat flow downstream of the active jet. In order to simplify the calculation, only the sheet model containing three nozzles was simulated, and periodic boundary conditions were used on both sides. At this time, the width of the inlet duct model was only 7.5mm. Figure 2 It shows the flow vortex structure generated by the interaction between the jet flow and the incoming flow, and the entire process of the flow vortex structure gradually becoming unstable, breaking up, and finally transitioning downstream. Figure 3 The wall heat flux distribution is displayed. From the comparison of the wall heat flux with the full laminar flow calculation results, the full turbulence calculation results and the calculation results without control measures, it can be seen that the forced transition method obtained by the present invention effectively promotes the boundary layer transition position to move forward from the downstream of the second compression corner to the vicinity of the first compression corner, indicating the effectiveness of the present invention.

[0052] The above embodiments are merely illustrative of the principles and effects of the present invention, and are not intended to limit the present invention. Anyone familiar with the art may modify or alter the above embodiments without departing from the spirit and scope of the present invention. Therefore, all equivalent modifications or alterations made by a person of ordinary skill in the art without departing from the spirit and technical ideas disclosed by the present invention shall still be covered by the claims of the present invention.

Claims

1. An active jet control method for forced transition of boundary layer in an ultra-high speed inlet, characterized in that: A row of equally spaced spray holes is arranged on the front body or the compression surface of the air intake of the air-breathing hyperspeed aircraft in the direction perpendicular to the flow direction. The gas source is ejected from the spray holes through the delivery pipeline. The ejected gas interacts with the incoming air to generate a flow vortex structure, inducing the hyperspeed boundary layer to forcibly transition to turbulence. The control method includes determining the installation position of the spray hole, determining the spray medium, determining the diameter of the spray hole, determining the total pressure before spraying, determining the spray hole spacing and the number of spray holes, and includes the following steps: (1) Determine the nozzle installation position. The nozzle is installed on the front body or the compression surface of the air intake of the air-breathing hypersonic vehicle; For multi-wave compression inlet, the nozzle is installed on the first compression surface. For the three-dimensional curved compression inlet, the nozzle is installed on the front body to ensure that the streamwise vortex induced by the jet flow has enough instability distance; At the same time, the nozzle installation position also needs to consider whether there is enough space in the aircraft to place the gas pipeline. After determining the installation position, the laminar boundary layer thickness δ at the installation position is obtained through fluid mechanics calculation in the laminar state. Here, the boundary layer thickness δ is defined as the total enthalpy boundary layer thickness. The total enthalpy at the outer edge of the boundary layer is 99% of the total enthalpy of the incoming flow. (2) Determine the injection medium. The injection medium shall be air or nitrogen, and the gas storage pressure shall not be less than 1 atmosphere; (3) Determine the nozzle diameter of 0.3-1mm and the total pressure before spraying P j The nozzle shape is circular and perpendicular to the compression surface, so that the jet airflow is perpendicular to the airflow in the boundary layer; the airflow penetration depth h It is taken as 0.5 times the thickness of the local boundary layer. The airflow penetration depth is defined as: Get the flow field Mach number distribution on the center line of the vertical nozzle, and define the distance from the point on the center line of the vertical nozzle when the Mach number first reaches the local maximum to the wall as the penetration depth h , since the air flow penetration depth satisfies the relationship: ,in p 2 is the static pressure of the sonic jet airflow, p is the local fluid static pressure when there is no injection, and the parameter a , b Calibration is performed through fluid mechanics calculations to obtain parameters a , b Then, substitute into the relational equation to obtain the static pressure of the jet flow p 2. According to the isentropic relationship ,in γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ; (4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is based on the penetration depth. h The number of installed nozzles is determined by the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The width w of the fluid entering the inlet duct is obtained through the wall limit streamlines. The width w at the nozzle installation position is the effective width range of the nozzle. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w / d.

2. The method for controlling active jetting in a super-high-speed inlet boundary layer forced transition according to claim 1, characterized in that: If there is already a nitrogen tank in the aircraft, there is no need for an additional gas tank. Just introduce a pipe from the nitrogen tank to connect to the spray hole, and then control it through the solenoid valve and the pressure reducing valve.

3. The method for controlling active jetting in a super-high-speed inlet boundary layer forced transition according to claim 1, characterized in that: If there is no existing nitrogen gas source in the aircraft, air is used as the injection medium, and an additional gas tank is equipped, and the gas storage pressure of the gas tank is not less than 1 atmosphere.

4. The method for controlling active jetting in a super-high-speed inlet boundary layer forced transition according to claim 1, characterized in that: The nozzle diameter is 0.5mm.

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