Widely adaptable variable geometry turbine and flow regulation structure for variable cycle aircraft engines

Abstract

The present invention discloses a variable-cycle aircraft engine with wide adaptability and a variable-geometry turbine and a flow regulating structure thereof. The turbine guide vane in the structure is integrally arranged at the upstream position of the turbine rotor and has a plurality of guide vane blades. A front cavity and a rear cavity are provided on each blade. The cavity is connected to the last stage of the compressor through an air bleed interface to introduce high-pressure airflow. The front airflow nozzle of the suction surface and the rear airflow nozzle of the pressure surface are used to spray the airflow into the mainstream channel, forming an aerodynamic blockage effect of local flow velocity reduction, and adjusting the throat area and airflow angle characteristics of the guide vane. The airflow regulating device includes a flow regulating valve and a pressure regulating valve to control the flow and pressure of the airflow, accurately adjust the airflow injection characteristics, and optimize the turbine flow and efficiency. The invention can effectively improve the turbine flow regulation accuracy, reduce mechanical complexity, improve turbine efficiency and engine life, and adapt to the needs under different flight conditions, and has important application value.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine turbine propulsion technology, specifically turbine flow regulation technology. Specifically, it relates to a variable-geometry turbine with wide adaptability for variable-cycle aero-engines and its flow regulation structure. By fabricating airflow nozzles on the suction and pressure surfaces of the turbine guide vane, and utilizing compressor bleed air to adjust the airflow ejected from the nozzles, the turbine guide vane throat area and the inlet flow angle of the turbine rotor are adjusted, thereby adjusting the turbine flow rate and the operating conditions of the turbine rotor. This structure is suitable for adaptive variable-cycle engines, allowing the engine to adjust the flow rate entering the turbine without the need for complex mechanical adjustment mechanisms. This simultaneously improves turbine efficiency, enabling wide flow range regulation and wide high-efficiency range operation, thereby reducing engine fuel consumption. Background Art

[0002] The Adaptive Cycle Engine (ACE) is considered the next major advancement in jet engine technology. Its core advantage lies in its ability to adjust the engine's flow rate, pressure ratio, and bypass ratio based on mission requirements by varying the geometry, size, or position of certain engine components. This allows for different thermodynamic cycles and optimizes engine performance across the entire flight envelope. During subsonic flight, the ACE operates in turbofan mode, offering low noise and low fuel consumption. During supersonic cruise, the ACE operates in turbojet mode, offering high thrust and maneuverability. This requires aircraft engines to achieve both low fuel consumption during subsonic flight and high thrust during supersonic flight, characteristics that traditional fixed-geometry engines struggle to meet. ACE engines, however, demonstrate significant potential in this area.

[0003] As a key component of an aircraft engine, the turbine's design and performance directly impact engine efficiency. For adaptive variable-cycle engines, a turbine with a variable-area turbine (VAT) is a core component. The application of VAT technology enables the turbine to maintain optimal efficiency over a wide operating range. VAT technology adjusts the turbine's power output by adjusting its flow capacity to match the fan or compressor's operating conditions. The adjustable range of VAT technology largely determines the adjustable range of the variable-cycle engine's bypass ratio, which in turn affects the engine's overall performance. Specifically, as flight speed and altitude change, the variable-geometry turbine automatically adjusts its geometry based on airflow conditions. By varying the angle of the turbine's adjustable guide vanes, the guide vane throat area and the rotor inlet airflow angle are adjusted, thereby controlling the turbine's flow rate and power output, ensuring the engine always operates at optimal levels. This not only improves overall engine performance but also extends its service life.

[0004] Currently, the primary method for implementing variable-geometry turbines is to use a mechanical adjustment mechanism to rotate the guide vanes. While varying the turbine's throat area through variable geometry effectively alters the turbine's geometry, it also suffers from inherent drawbacks and limitations. First, traditional mechanical variable-geometry adjustment mechanisms suffer from structural complexity, increased weight, and reduced reliability. To prevent collision and friction during guide vane rotation, a certain gap remains between the vanes, the hub, and the casing. This results in "excessive losses" such as leakage losses at the guide vane end and blade angle of attack losses. When the guide vane adjustment angle is too large, these "excessive losses" increase significantly, severely impacting turbine efficiency and negating the benefits of variable turbine geometry. Second, traditional variable-geometry turbine design approaches struggle to reconcile the flow rate adjustment range and the high-efficiency operating range, negatively impacting the engine's wide-range, high-efficiency operation. Furthermore, excessively large guide vane adjustment angles complicate the design of the adjustable mechanism. Mechanical adjustment mechanisms also suffer from slow response and high technical risks associated with high-temperature operation.

[0005] In summary, traditional variable geometry turbines have many technical difficulties in flow regulation, such as complex mechanical structure, large aerodynamic losses, difficulty in maintaining efficiency, and slow response speed. Therefore, how to develop a variable geometry turbine flow regulation structure with simple structure, fast response speed, high efficiency, strong reliability and wide adaptability suitable for adaptive variable cycle engines is a technical problem that needs to be solved urgently. Summary of the Invention

[0006] (1) Purpose of the invention

[0007] Changing the fixed guide vane to an adjustable one increases the design difficulty of the adjustable mechanism, significantly increasing the weight and cost of the turbine. Furthermore, repeated adjustment over long periods of time can lead to blade wear and fatigue, shortening the engine life. While adjustable guide vanes offer significant advantages in flow control and can improve engine performance under different operating conditions, their inherent drawbacks and limitations cannot be ignored. To address these and other inherent drawbacks of the variable geometry turbine, including increased weight due to a complex mechanical structure, low reliability of the adjustment mechanism, large leakage losses in the end zone, significant reduction in aerodynamic efficiency, slow adjustment response, and high technical risks in high-temperature environments, and to address at least one of the aforementioned and other technical issues in the prior art, the present invention provides a variable geometry turbine with wide adaptability for a variable cycle aero-engine and a flow control structure thereof. Specifically, the variable geometry turbine flow control structure for a variable cycle aero-engine with wide adaptability comprises a flow control structure that utilizes airflow nozzles formed on the suction and pressure surfaces of the turbine guide vane, utilizes high-pressure airflow from the final stage of the compressor, and injects it into the turbine guide vane through cavities formed inside the vanes, creating a blockage on the suction surface of the guide vane. This structure then adjusts the turbine guide vane throat area and the turbine rotor inlet airflow angle, achieving precise regulation of the turbine flow and operating conditions. This structure is different from the turbine cooling airflow. The aerodynamic adjustment will not increase the complexity of the mechanical structure, avoiding the wear and fatigue problems of traditional adjustable guide vanes in high temperature environments. It can effectively regulate the turbine flow and the moving blade inlet angle, and can maintain a high turbine efficiency in a wide range of operating conditions. At the same time, it has the advantages of simple structure, light weight, fast response, and high reliability, and is very suitable for adaptive variable cycle engines.

[0008] (2) Technical solution

[0009] In order to achieve the purpose of the invention and solve the technical problems, the present invention adopts the following technical solutions:

[0010] The first object of the present invention is to provide a variable-cycle aircraft engine with wide adaptability and variable-geometry turbine flow control structure, which is used to adjust the turbine guide vane throat area and the turbine rotor inlet airflow angle, thereby achieving precise adjustment of turbine flow and optimization of turbine efficiency under different flight conditions. The structure comprises at least a turbine guide vane, a compressor bleed air channel, and an airflow control device. Specifically:

[0011] The turbine guide vane is integrally arranged at the upstream position of the turbine rotor, and includes an outer casing, an inner casing, and a plurality of guide vanes uniformly distributed circumferentially between the outer casing and the inner casing, and the space between each adjacent guide vane forms a turbine guide vane main flow channel, each of the guide vanes includes a blade suction surface, a blade pressure surface, and at least one closed front cavity and one closed rear cavity machined and arranged on its main body, wherein:

[0012] The front cavity and the rear cavity are both extended along the span direction of the blade and their overall height is equivalent to the height of the blade. The front cavity is arranged near the leading edge of the blade, and the rear cavity is arranged near the trailing edge of the blade. The bottoms of the front cavity and the rear cavity are both processed and provided with bleed air interfaces connected to the cavities. Each of the bleed air interfaces is connected to the last stage of the compressor through the compressor bleed air channel, and is used to guide the high-pressure airflow of the last stage of the compressor into the front cavity and the rear cavity of the guide vane;

[0013] A plurality of suction front airflow nozzles distributed along the blade span direction are machined and provided at an upstream portion of the blade suction surface adjacent to the blade leading edge, and each of the suction front airflow nozzles is communicated with the front cavity, and is used to inject the compressor final stage high-pressure airflow in the front cavity into the main flow channel of the turbine guide vane, and form an aerodynamic blockage effect of local flow velocity reduction on the blade suction surface, thereby adjusting the effective throat area and airflow angle characteristics of the main flow channel of the turbine guide vane;

[0014] The downstream portion of the pressure surface of the blade between the trailing edge of the blade and the throat is processed and provided with a plurality of pressure surface tail airflow nozzles distributed along the span direction of the blade, and each of the pressure surface tail airflow nozzles is connected to the rear cavity, and is used to inject the high-pressure airflow of the last stage of the compressor in the rear cavity into the main flow channel of the turbine guide vane and change the airflow distribution characteristics of the tail of the pressure surface of the blade, adjust the airflow velocity and flow rate of the tail of the pressure surface of the blade, increase the guiding effect of the tail airflow, and thus optimize the outlet airflow angle of the main flow channel of the turbine guide vane and the inlet airflow angle of the turbine rotor;

[0015] The airflow regulating device includes at least a flow regulating valve and a pressure regulating valve arranged on the compressor bleed air channel, the flow regulating valve is used to regulate the high-pressure bleed air flow entering the corresponding front cavity and rear cavity, and the pressure regulating valve is used to regulate the high-pressure bleed air pressure entering the corresponding front cavity and rear cavity, and the flow regulating valve and the pressure regulating valve are both communicated with the engine control system to accurately control the injection airflow characteristics of the front airflow nozzle of the suction surface and the rear airflow nozzle of the pressure surface according to different flight conditions and engine working conditions, so as to realize dynamic adjustment of the throat area of ​​the turbine guide vane mainstream channel, the outlet airflow angle and the turbine rotor inlet airflow angle.

[0016] The second object of the present invention is to provide a wide-adaptability variable geometry turbine, including the above-mentioned variable-cycle aviation engine wide-adaptability variable geometry turbine flow regulating structure.

[0017] The third object of the present invention is to provide an adaptive variable cycle engine, including the above-mentioned wide adaptability variable geometry turbine.

[0018] (3) Technical effects

[0019] Compared with the prior art, the variable-cycle aircraft engine wide-adaptability variable-geometry turbine and its flow regulation structure of the present invention have the following beneficial and significant technical effects:

[0020] (1) The present invention adopts a turbine flow control method based on the principle of aerodynamic regulation, eliminating the need for complex mechanical adjustment mechanisms and avoiding the problems of increased weight, reduced reliability, and high maintenance costs associated with the mechanical structure of traditional variable geometry turbines. By utilizing compressor bleed air, aerodynamic blocking and guiding effects are formed on the suction and pressure surfaces of the guide vanes, achieving precise control of the turbine flow, simplifying the structure, improving reliability, and reducing maintenance costs.

[0021] (2) The present invention can flexibly adjust the guide vane throat area and the rotor inlet airflow angle by controlling the flow rate and pressure of the compressor bleed air, thereby achieving precise control of the turbine flow rate within a wider range of flight conditions. The response speed of the aerodynamic adjustment method is much higher than that of the traditional mechanical adjustment method, which can better meet the needs of the adaptive variable cycle engine for rapid adjustment and improve the engine's adaptability to changes in flight conditions.

[0022] (3) The present invention effectively reduces the "extra loss" of the turbine. During the guide vane adjustment process, the traditional variable geometry turbine will generate additional leakage losses due to the gap between the blades and the hub and casing. The present invention changes the turbine flow rate through aerodynamic adjustment, avoiding the large rotation of the guide vanes, thereby effectively reducing these "extra losses", improving the efficiency of the turbine, and thus reducing the fuel consumption of the engine. In addition, by arranging nozzles on the suction side and the pressure side respectively, the present invention can more finely control the guide throat area and the rotor inlet airflow angle, thereby optimizing the aerodynamic performance of the turbine under different working conditions, improving the efficiency of the turbine, and widening the efficient working range of the turbine. This enables the engine to maintain optimal performance within a wider flight envelope, improving the overall performance of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] The accompanying drawings, which constitute part of the present invention, are provided to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are provided to explain the present invention and do not constitute undue limitations thereon. The embodiments of the present invention will be described in detail below with reference to the accompanying drawings, wherein:

[0024] Figure 1 The figure shows a schematic diagram of the variable geometry turbine flow rate regulation structure of a variable cycle aviation engine with wide adaptability according to the present invention;

[0025] Figure 2 Shown is a meridian view of a turbine guide vane with aerodynamic adjustment according to the present invention, wherein (a) is a spanwise slot-type jet structure and (b) is a spanwise hole-type jet structure.

[0026] Explanation of reference numerals: 1 is the outer casing; 2 is the airflow nozzle at the front of the suction surface, 2-1 is the air jet slit at the front of the suction surface, and 2-2 is the air jet hole at the front of the suction surface; 3 is the inner casing; 4 is the guide vane; 5 is the rear cavity of the blade; 6 is the front air bleed interface inside the blade; 7 is the rear air bleed interface inside the blade; 8 is the front cavity of the blade; 9 is the suction surface of the blade; 10 is the pressure surface of the blade; 11 is the front throat of the aerodynamic adjustment; 12 is the rear throat of the aerodynamic adjustment; 13 is the grid pitch; 14 is the airflow nozzle at the rear of the pressure surface, 14-1 is the air jet slit at the rear of the pressure surface, and 14-2 is the air jet hole at the rear of the pressure surface; L is the distance between the air jet slit and the leading edge, which is usually 10%-20% of the axial chord length; d is the nozzle width; α 2, α 3 are the outlet airflow angles before and after pneumatic adjustment respectively. DETAILED DESCRIPTION

[0027] In order to make the purpose, technical solutions and advantages of the implementation of the present invention clearer, the technical solutions in the embodiments of the present invention will be described in more detail below with reference to the drawings in the embodiments of the present invention. The described embodiments are part of the embodiments of the present invention, rather than all the embodiments. The embodiments described below with reference to the drawings are exemplary and are intended to be used to explain the present invention, and should not be understood as limiting the present invention. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0028] The present invention aims to provide a variable-geometry turbine with wide adaptability for a variable-cycle aircraft engine and its flow regulation structure. By adding an adjustable nozzle upstream of the throat of the turbine guide vane suction surface, the nozzle utilizes high-pressure airflow from the final compressor stage. By machining cavities within the blades, the nozzle injects the main flow into the turbine guide vane at different locations and angles, creating a blockage on the guide vane suction surface. This allows the turbine guide vane throat area and the turbine rotor inlet airflow angle to be adjusted, achieving precise regulation of turbine flow and operating conditions. Unlike turbine cooling airflow, this structure's aerodynamic regulation does not increase the complexity of the mechanical structure, avoiding the wear and fatigue issues associated with conventional adjustable guide vanes in high-temperature environments. Furthermore, it can effectively regulate turbine flow and blade inlet angle, maintaining high turbine efficiency over a wide range of operating conditions.

[0029] As a specific example, Figure 1 、 2As shown, the variable-cycle aviation engine wide-adaptability variable-geometry turbine flow regulation structure of the present invention mainly includes a turbine guide vane, a compressor bleed air channel and an airflow regulation device. The core component of this structure is the turbine guide vane, which is arranged as a whole at the upstream position of the turbine rotor, including an outer casing 1, an inner casing 3 and a plurality of guide vanes 4 uniformly distributed along the circumference between the outer casing 2 and the inner casing 3, and the space between each adjacent guide vane 4 forms a turbine guide vane main flow channel. Each guide vane 4 includes a blade suction surface 9 and a blade pressure surface 10, and at least one closed front cavity 8 and a closed rear cavity 5 are machined and arranged on the main body of the blade. This structure accurately adjusts the turbine flow and the rotor inlet airflow angle through the airflow regulation device, thereby optimizing the turbine efficiency and improving the comprehensive performance of the engine under different flight conditions.

[0030] More specifically, the front cavity 8 and the rear cavity 5 both extend along the span direction of the blade, and their heights are substantially the same as the blade height, to ensure that the bleed air flow can be evenly distributed over the entire blade height. The front cavity 8 is arranged close to the leading edge of the blade, and is used to provide a high-pressure air source for the front airflow nozzle 2 of the suction surface; the rear cavity 5 is arranged close to the trailing edge of the blade, and is used to provide a high-pressure air source for the rear airflow nozzle 14 of the pressure surface. A front bleed air interface 6 is provided at the bottom of the front cavity 8, and a rear bleed air interface 7 is provided at the bottom of the rear cavity 5. These bleed air interfaces are connected to the last stage of the compressor through the compressor bleed air channel, and the high-pressure airflow of the last stage of the compressor is guided to the front cavity 8 and the rear cavity 5, respectively. Since the airflow pressure of the last stage of the compressor is relatively high, it can provide sufficient energy to drive the jet airflow, thereby effectively adjusting the airflow characteristics of the turbine guide vane.

[0031] At the upstream position of the blade suction surface 9 near the leading edge, a plurality of suction front airflow nozzles 2 are machined and arranged along the blade span. Each suction front airflow nozzle 2 is connected to the front cavity 8 and is used to inject the high-pressure airflow of the last stage of the compressor in the front cavity 8 into the main flow channel of the turbine guide vane. The injected airflow forms an aerodynamic blockage effect on the blade suction surface 9, which reduces the local flow velocity, which is equivalent to reducing the effective throat area of ​​the main flow channel of the turbine guide vane. Figure 1 As shown, 11 is the throat before pneumatic adjustment; 12 is the throat after pneumatic adjustment. At the same time, the aerodynamic blockage effect also changes the airflow direction, affecting the airflow angle characteristics of the turbine guide vane main channel, such as Figure 1 As shown in α 2, α3 represent the outlet airflow angles before and after aerodynamic adjustment. By adjusting the pressure and flow rate of the jet airflow, the effective throat area and airflow angle can be precisely controlled, thereby achieving precise regulation of the turbine flow rate. Furthermore, the jet angle and flow rate of the airflow nozzle 2 at the front of the suction face can be dynamically optimized according to different flight conditions (such as low speed, high speed, cruising, and acceleration). By controlling the pressure, flow rate, and angle of the jet airflow, the flow characteristics of the turbine guide vane's main channel can be precisely adjusted, optimizing turbine flow rate and efficiency. This also effectively reduces airflow disturbances and minimizes energy loss during turbine flow adjustment.

[0032] On the downstream position between the trailing edge and the throat of the blade pressure surface 10, a plurality of pressure surface tail airflow nozzles 14 distributed along the blade span are machined and provided. Each pressure surface tail airflow nozzle 14 is connected to the rear cavity 5 and is used to inject the high-pressure airflow of the last stage of the compressor in the rear cavity 5 into the main channel of the turbine guide vane. The injected airflow changes the airflow distribution characteristics at the tail of the blade pressure surface 10, adjusts the airflow velocity and flow at the tail of the blade pressure surface 10, increases the guiding effect of the tail airflow, and thus optimizes the outlet airflow angle of the main channel of the turbine guide vane and the inlet airflow angle of the turbine rotor. In addition, the injection angle and flow rate of the pressure surface tail airflow nozzle 14 can be dynamically adjusted according to the flight conditions to optimize the guiding effect of the tail airflow and effectively reduce the instability of the airflow, thereby improving the turbine efficiency. By working in conjunction with the front airflow nozzle of the suction surface, the pressure surface tail airflow nozzle can form a uniform airflow distribution in the main channel of the turbine guide vane, reducing the airflow fluctuations during the turbine flow regulation process.

[0033] The airflow regulating device of the present invention includes at least a flow regulating valve and a pressure regulating valve provided on the compressor bleed air passage. The flow regulating valve is used to regulate the flow rate of high-pressure bleed air entering the corresponding front cavity 8 and rear cavity 5, and the pressure regulating valve is used to regulate the pressure of high-pressure bleed air entering the corresponding front cavity 8 and rear cavity 5. Both the flow regulating valve and the pressure regulating valve are communicatively connected to the engine control system. The engine control system precisely controls the jet flow characteristics of the front airflow nozzle 2 of the suction surface and the rear airflow nozzle 14 of the pressure surface according to different flight conditions and engine operating states, thereby achieving dynamic regulation of the throat area of ​​the turbine guide vane mainstream passage, the outlet airflow angle, and the turbine rotor inlet airflow angle.

[0034] As a preferred solution, the arrangement position of the suction front front airflow nozzle 2 on the suction surface 9 of the blade is such that the distance L from the leading edge of the blade in the axial direction is 10% to 20% of the axial chord length of the blade. Figure 1The nozzle structure can be either a spanwise slot nozzle 2-1 or a spanwise discrete hole nozzle 2-2. The spanwise slot nozzle 2-1 consists of jet slots arranged continuously along the blade span, while the spanwise discrete hole nozzle 2-2 consists of a number of jet holes spaced apart along the blade span. The jet slot width d and the jet hole diameter d are 1% to 5% of the blade axial chord length to ensure a uniform and stable aerodynamic blocking effect without compromising blade strength.

[0035] As a preferred option, the pressure surface aft airflow nozzle 14 is positioned in the downstream region between the blade trailing edge and the throat. The nozzle structure can also be either a spanwise slot nozzle 14-1 or a spanwise discrete hole nozzle 14-2. Its spray direction forms an angle of 10° to 30° with the surface normal of the blade pressure surface 10 to optimize airflow guidance and form a continuous spanwise air film aft of the blade pressure surface 10, uniformly changing the airflow distribution characteristics aft of the blade pressure surface and optimizing the turbine rotor inlet airflow angle.

[0036] As a preferred embodiment, the front cavity 8 and the rear cavity 5 are both independent, sealed hollow cavity structures, whose height is consistent with the span-wise height of the guide vane 4. Their cross-sectional shapes can be circular, elliptical, or rectangular. The axial width of the front cavity 8 is 15% to 25% of the axial chord length of the blade, and the axial width of the rear cavity 5 is 20% to 30% of the axial chord length of the blade. This design ensures that the cavity has sufficient structural strength and durability under the action of high-pressure airflow, while reducing the weakening effect of the cavity on the internal structure of the blade and ensuring uniform distribution of the jet airflow along the height direction of the blade.

[0037] As a preferred solution, the compressor bleed air channel includes a main bleed air duct and multiple branch bleed air ducts. The main bleed air duct is connected to the last stage of the compressor, and the branch bleed air ducts are respectively connected to the bleed air interfaces of the front cavity 8 and the rear cavity 5 of the corresponding guide vane 4. An independent flow regulating valve and a pressure regulating valve are provided on each branch bleed air duct to realize independent control of the jet airflow of each blade, so that the turbine flow can be adjusted more finely. The flow regulating valve and the pressure regulating valve in the airflow regulating device are both electronic control valves, and each electronic control valve is communicatively connected to the engine control system. The engine control system dynamically adjusts the opening of the flow regulating valve and the pressure regulating valve according to the real-time operating parameters of the engine to accurately control the high-pressure airflow characteristics of the front cavity 8 and the rear cavity 5, thereby realizing precise control of the turbine flow and optimization of the turbine efficiency under different flight conditions.

[0038] Reference Figure 1 The working principle of the variable-cycle aviation engine wide-adaptability variable-geometry turbine flow regulation structure of the present invention is as follows:

[0039] First, the high-pressure airflow from the final stage of the compressor passes through the compressor bleed air duct and enters the enclosed front cavity 8 and enclosed rear cavity 5 inside the turbine guide vane 4. The enclosed front cavity 8 and rear cavity 5 are located near the leading and trailing edges of the turbine guide vane 4, respectively, and extend along the span of the blade to ensure uniform distribution of the bleed airflow. Bleed air ports 6 and 7 are located at the bottom of the front cavity 8 and rear cavity 14, respectively, and are connected to the compressor bleed air duct.

[0040] Secondly, the bleed air in the front cavity 8 is ejected into the main flow channel of the turbine guide vane through the suction face front airflow nozzle 2 located upstream of the leading edge of the blade suction face 9. The ejected airflow forms a local area of ​​reduced flow velocity on the blade suction face 9, producing an aerodynamic blocking effect. This blocking effect is equivalent to reducing the throat area of ​​the turbine guide vane (such as Figure 1 As shown, 11 is the throat before pneumatic adjustment, and 12 is the throat after pneumatic adjustment. It can be seen that the throat area is reduced through pneumatic adjustment), and the airflow direction is changed, which affects the turbine rotor inlet airflow angle ( Figure 1 The outlet airflow angle before pneumatic adjustment α 2Change to the outlet airflow angle after aerodynamic adjustment α 3. Realize the control of the turbine rotor inlet airflow angle).

[0041] At the same time, the bleed air in the rear cavity 5 is ejected into the main flow channel of the turbine guide vane through the pressure surface tail airflow nozzle 14 located between the trailing edge of the blade pressure surface 10 and the throat. The ejected airflow changes the airflow distribution characteristics at the tail of the blade pressure surface 10, adjusts the airflow velocity and flow at the tail of the blade pressure surface 10, increases the guiding effect of the tail airflow, and further optimizes the turbine rotor inlet airflow angle.

[0042] Finally, the flow rate and pressure of the bleed air entering the front cavity 8 and the rear cavity 5 are precisely controlled by the air flow regulating device. The air flow regulating device includes a flow regulating valve and a pressure regulating valve arranged on the compressor bleed air channel. The flow regulating valve is used to regulate the bleed air flow entering the corresponding front cavity 8 and the rear cavity 5, and the pressure regulating valve is used to regulate the bleed air pressure entering the corresponding front cavity 8 and the rear cavity 5. These valves are communicated with the engine control system. The engine control system adjusts the opening of the flow regulating valve and the pressure regulating valve in real time according to different flight conditions and engine working conditions, such as flight speed, altitude, engine speed, thrust requirement, etc., so as to precisely control the injection airflow characteristics of the front airflow nozzle 2 of the suction surface and the rear airflow nozzle 14 of the pressure surface, realize dynamic adjustment of the throat area and the turbine rotor inlet airflow angle, and finally realize precise adjustment of the turbine flow and optimization of turbine efficiency under different flight conditions.

[0043] The above embodiments fully and effectively achieve the objectives of the present invention. Those skilled in the art will appreciate that the present invention includes, but is not limited to, the contents described in the accompanying drawings and the above specific embodiments. Although the present invention has been described with reference to the embodiments currently considered to be the most practical and preferred, it should be understood that the present invention is not limited to the disclosed embodiments, and any modifications that do not deviate from the functional and structural principles of the present invention are intended to be included within the scope of the claims.

Claims

1. A variable-cycle aircraft engine with wide adaptability and variable-geometry turbine flow control structure, comprising at least a turbine guide vane, a compressor bleed air passage, and an airflow control device, characterized in that: The turbine guide vane is integrally arranged at the upstream position of the turbine rotor, and includes an outer casing, an inner casing, and a plurality of guide vanes uniformly distributed circumferentially between the outer casing and the inner casing, and the space between each adjacent guide vane forms a turbine guide vane main flow channel, each of the guide vanes includes a blade suction surface, a blade pressure surface, and at least one closed front cavity and one closed rear cavity machined and arranged on its main body, wherein: The front cavity and the rear cavity are extended along the span direction of the blade as a whole and their overall height is equivalent to the height of the blade. The front cavity is arranged near the leading edge of the blade, and the rear cavity is arranged near the trailing edge of the blade. The bottoms of the front cavity and the rear cavity are both processed and provided with bleed air interfaces connected to the cavities. Each of the bleed air interfaces is connected to the last stage of the compressor through the compressor bleed air channel, and is used to guide the high-pressure airflow of the last stage of the compressor into the front cavity and the rear cavity of the guide vane; A plurality of suction front airflow nozzles distributed along the blade span direction are machined and provided at an upstream portion of the blade suction surface adjacent to the blade leading edge, and each of the suction front airflow nozzles is communicated with the front cavity, and is used to inject the compressor final stage high-pressure airflow in the front cavity into the main flow channel of the turbine guide vane, and form an aerodynamic blockage effect of local flow velocity reduction on the blade suction surface, thereby adjusting the effective throat area and airflow angle characteristics of the main flow channel of the turbine guide vane; The downstream portion of the pressure surface of the blade between the trailing edge of the blade and the throat is processed and provided with a plurality of pressure surface tail airflow nozzles distributed along the span direction of the blade, and each of the pressure surface tail airflow nozzles is connected to the rear cavity, and is used to inject the high-pressure airflow of the last stage of the compressor in the rear cavity into the main flow channel of the turbine guide vane and change the airflow distribution characteristics of the tail of the pressure surface of the blade, adjust the airflow velocity and flow rate of the tail of the pressure surface of the blade, increase the guiding effect of the tail airflow, and thus optimize the outlet airflow angle of the main flow channel of the turbine guide vane and the inlet airflow angle of the turbine rotor; The airflow regulating device includes at least a flow regulating valve and a pressure regulating valve arranged on the compressor bleed air channel, the flow regulating valve is used to regulate the high-pressure bleed air flow entering the corresponding front cavity and rear cavity, and the pressure regulating valve is used to regulate the high-pressure bleed air pressure entering the corresponding front cavity and rear cavity, and the flow regulating valve and the pressure regulating valve are both communicated with the engine control system to accurately control the injection airflow characteristics of the front airflow nozzle of the suction surface and the rear airflow nozzle of the pressure surface according to different flight conditions and engine working conditions, so as to realize dynamic adjustment of the throat area of ​​the turbine guide vane mainstream channel, the outlet airflow angle and the turbine rotor inlet airflow angle.

2. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1 is characterized in that: The arrangement position of the airflow nozzle in front of the suction surface on the suction surface of the blade is 10%-20% of the axial chord length of the blade from the leading edge of the blade in the axial direction. The nozzle structure is a span-wise slot nozzle or a span-wise discrete hole nozzle, wherein the span-wise slot nozzle is an air jet slot continuously arranged along the span direction of the blade, and the span-wise discrete hole nozzle is a plurality of air jet holes arranged at intervals along the span direction of the blade, and the width of the air jet slot and the diameter of the air jet hole are 1-5% of the axial chord length of the blade.

3. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1, characterized in that: The arrangement position of the airflow nozzle at the tail of the pressure surface is located in the downstream area between the trailing edge of the blade and the throat. The nozzle structure is one of a span-wise slot nozzle or a span-wise discrete hole nozzle, and the injection direction is at an angle of 10°-30° to the surface normal of the blade pressure surface.

4. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1, characterized in that: The size, number and injection angle of the airflow nozzles in front of the suction surface are optimized according to the required aerodynamic blockage effect, and the size, number and injection angle of the airflow nozzles in the rear of the pressure surface are optimized according to the required airflow guiding effect. By comprehensively considering the coordinated work of the suction surface and pressure surface nozzles, the optimal combination of nozzle size, number and injection angle is finally determined to achieve precise adjustment of turbine flow and optimization of turbine efficiency under different flight conditions.

5. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1, characterized in that: The aerodynamic blocking effect and guiding effect formed by the jet airflow of the suction surface front airflow nozzle and the pressure surface rear airflow nozzle respectively cover the throat area and outlet area of ​​the turbine guide vane main flow channel, and the intensity of the aerodynamic blocking effect is optimized by adjusting the pressure and flow rate of the jet airflow to adapt to the airflow characteristic requirements under different flight conditions.

6. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1, characterized in that: The front cavity and the rear cavity are both independent and closed hollow cavity structures, and their heights are consistent with the span-wise height of the guide vane. Their cross-sectional shapes are circular, elliptical or rectangular. The axial width of the front cavity is 15%-25% of the axial chord length of the blade, and the axial width of the rear cavity is 20%-30% of the axial chord length of the blade.

7. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1, characterized in that: The compressor bleed air channel includes a main bleed air duct and multiple branch bleed air ducts. The main bleed air duct is connected to the last stage of the compressor, and each branch bleed air duct is respectively connected to the bleed air interface of the corresponding front cavity and rear cavity, and each branch bleed air duct is provided with an independent flow regulating valve and pressure regulating valve to realize independent control of the jet airflow of each blade.

8. The variable cycle aircraft engine wide adaptability variable geometry turbine flow regulating structure according to claim 1, characterized in that: The flow regulating valve and pressure regulating valve in the airflow regulating device are both electronically controlled valves, and each electronically controlled valve is communicatively connected to the engine control system. The engine control system dynamically adjusts the opening of the flow regulating valve and the pressure regulating valve according to the real-time operating parameters of the engine to accurately control the high-pressure airflow characteristics of the front cavity and the rear cavity.

9. A variable geometry turbine with wide adaptability, characterized in that: It includes the variable-cycle aviation engine wide-adaptability variable-geometry turbine flow regulation structure as described in any one of claims 1 to 8.

10. An adaptive variable cycle engine, characterized in that: Including the wide adaptability variable geometry turbine as described in claim 9.

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