A blade or airfoil with built-in adaptive vortex generators

Abstract

The application discloses a blade or wing with built-in self-adaptive vortex generator, comprising a compressor blade or aircraft wing, a static pressure cavity, a pressure surface static pressure slit, a suction surface static pressure slit, a sliding block, a vortex generator, a spring cavity and a spring; the static pressure cavity is a hollow structure in the compressor blade or aircraft wing; the pressure surface static pressure slit is located on the pressure surface of the compressor blade or aircraft wing; the suction surface static pressure slit is located on the suction surface of the compressor blade or aircraft wing; the static pressure cavity is built-in with a movable sliding block; the side of the sliding block, which is opposite to the suction surface static pressure slit, is installed with the vortex generator; the vortex generator can extend to the outside of the compressor blade or aircraft wing through the suction surface static pressure slit; the static pressure cavity is connected with the hollow spring cavity; and the sliding block is connected with the spring cavity wall surface through the spring; the vortex generator can be automatically started and stopped according to different attack angle conditions without introducing external air source or power supply.

Description

Technical Field

[0001] This invention relates to a blade or wing with a built-in adaptive vortex generator, belonging to the technical field of compressors (internal flow) or aircraft wings (external flow). Background Technology

[0002] Improving the performance of compressors in the internal flow of aviation and wings in the external flow has been a long-term pursuit. Among various axial compressors used in applications such as aero gas turbine engines, achieving higher pressure ratios in single-stage compressors has been a long-standing development direction in the compressor field. Historically, increasing the stage pressure ratio of compressors has primarily been achieved through aerodynamic optimization, such as increasing blade rim velocity and airfoil design. However, the increase in blade rim velocity is limited by material strength limits, and aerodynamic optimization of airfoil design has limited impact on compressor performance at current design levels. When attempting to significantly increase the load on a single-stage compressor beyond current levels, flow separation on the blade back surface can occur, causing a sharp drop in compressor pressure ratio and efficiency, and even compressor stall. Furthermore, when an external flow wing is at a high angle of attack, its suction surface also experiences large-scale flow separation, leading to a sharp drop in the wing's lift-to-drag ratio and even stall. Therefore, various flow control technologies that can suppress or even eliminate flow separation have received considerable attention from researchers both domestically and internationally.

[0003] Based on existing work in the field of flow control both domestically and internationally, the relevant technologies can be mainly divided into two categories: (1) Passive flow control technologies, such as slotted blades, tandem blades, stirrups, eddy current generators, etc. (2) Active flow control technologies, such as adsorption compressors, synthetic jets, acoustic excitation, plasma excitation, traveling wave wall control, etc.

[0004] To meet the needs of future higher-load compressor blades and high-stall-margin airfoil designs, these technologies have significant shortcomings from a practical engineering application perspective. Passive flow control technology is characterized by not requiring external energy input and boasts advantages such as simple structure, ease of implementation, and long lifespan. However, its main drawback is that flow control parameters cannot be adjusted according to changes in operating conditions. For example, although slotted blades typically have lower flow losses than seamless blades under flow separation conditions, they exhibit greater flow losses in the absence of flow separation. Similarly, while mixed-flow typically has lower flow losses than non-mixed-flow under laminar upstream conditions, it may experience greater flow losses when the upstream flow transitions from laminar to turbulent. Active flow control technology requires external energy input. Its advantage lies in the ability to adjust flow control parameters according to changing operating conditions. However, it is typically complex in structure, requires an external air or power source, and necessitates numerous auxiliary devices, making its implementation difficult in practical engineering applications. For example, acoustic excitation devices often require speakers installed inside the blades, and also necessitate the introduction of additional power and circuit systems that occupy significant space and weight. Therefore, it is imperative to seek a new technology that combines the strong engineering practicality of passive flow control technology with the strong adaptability to operating conditions of active flow control technology. Summary of the Invention

[0005] The purpose of this invention is to achieve high aerodynamic performance for compressor blades or aircraft wings under multiple operating conditions. This ensures effective suppression of external flow separation at high angles of attack while avoiding additional shape drag and vortex drag at low angles of attack where separation is not possible. Specifically, a blade or wing with a built-in adaptive vortex generator is proposed. This generator can be automatically activated (including adjusting its release level) and deactivated according to different angles of attack without the need for an external air or power source.

[0006] To achieve the above objectives, the present invention is implemented using the following technical solution:

[0007] This invention provides a compressor blade or wing with a built-in adaptive vortex generator, comprising: a compressor blade or aircraft wing, a static pressure chamber, a pressure surface static pressure slot, a suction surface static pressure slot, a slider, a vortex generator, a spring chamber, and a spring; wherein, the static pressure chamber is a hollow structure within the compressor blade or aircraft wing, the pressure surface static pressure slot is located on the pressure surface of the compressor blade or aircraft wing, the suction surface static pressure slot is located on the suction surface of the compressor blade or aircraft wing, a movable slider is built into the static pressure chamber, and a vortex generator is mounted on the side of the slider facing the suction surface static pressure slot. The vortex generator can extend through the suction surface static pressure slot to the outside of the compressor blade or aircraft wing. The static pressure chamber and the hollow surface static pressure chamber are connected. The spring chambers are connected, and the slider is connected to the spring chamber wall through the spring. The static pressure seam on the pressure surface and the static pressure seam on the suction surface are respectively used to conduct the static pressure of the suction surface and pressure surface of the compressor blade or wing to the upper and lower surfaces of the slider. When the compressor blade or wing is at a small angle of attack and there is no suction surface flow separation, the pressure difference between the upper and lower sides of the slider is low. At this time, the vortex generator is completely inside the static pressure chamber and is in a non-working state. When the compressor blade or wing is at a large angle of attack and there is suction surface flow separation, the pressure difference between the upper and lower sides of the slider is high, causing the slider to move towards the suction surface, driving the vortex generator to extend out of the suction surface of the compressor blade or wing, so that the vortex generator is in a working state.

[0008] Furthermore, the height h of the vortex generator is 5% of the maximum thickness H of the compressor blade or aircraft wing. ～ 50%.

[0009] Furthermore, the height of the static pressure chamber H2 = H1 + h - H3, where H1 is the height of the slider and H3 is the height of the static pressure gap on the suction surface.

[0010] Furthermore, assuming that the compressor blades or aircraft wings are in the critical state of just separating and the maximum operating angle of attack, the pressure difference between the static pressure gap on the pressure surface and the static pressure gap on the suction surface is ΔP1 and ΔP2, respectively, where ΔP1 < ΔP2. When the actual pressure difference between the static pressure gap on the pressure surface and the static pressure gap on the suction surface is ΔP ≤ ΔP1, the lower surface of the slider is in close contact with the lower surface of the static pressure chamber, and the vortex generator is in an unextended state. When ΔP1 < ΔP < ΔP2, the lower surface of the slider leaves the lower surface of the static pressure chamber, and the vortex generator is in a partially extended state. When ΔP = ΔP2, the upper surface of the slider is in close contact with the upper surface of the static pressure chamber, and the vortex generator is in a fully extended state.

[0011] Furthermore, assuming the slider's spanwise thickness is w, when the eddy current generator is in a fully extended state, the spring preload F0 is calculated as follows: F0 = bwΔP1, where b is the width of the slider.

[0012] Furthermore, the spring constant k is calculated using the following formula: k = bw(△P2 - △P1) / (H2 - H1).

[0013] Compared with the prior art, the beneficial effects achieved by the present invention are as follows:

[0014] This invention provides a blade or wing with a built-in adaptive vortex generator. The vortex generator can be automatically activated and deactivated under different angles of attack conditions by relying solely on its own structure. This allows it to adaptively suppress external flow separation and reduce additional profile drag and vortex drag. It does not require an external air source or power source and has advantages such as simple structure, no need for an external energy source, strong adaptability, and strong engineering applicability. Attached Figure Description

[0015] Figure 1 This is a schematic diagram of a blade or wing with a built-in adaptive vortex generator (at a high angle of attack).

[0016] Figure 2 This is a schematic diagram of a blade or wing with a built-in adaptive vortex generator (at a small angle of attack).

[0017] Figure 3 This is a schematic diagram of an adaptive eddy current generator and slider (top view in the direction of the suction surface).

[0018] Figure 4 This is a magnified view of the adaptive eddy current generator in its fully "on" mode (i.e., at maximum angle of attack).

[0019] Figure 5 This is a magnified view of the adaptive eddy current generator in its completely "off" mode (small angle of attack).

[0020] In the diagram: 1. Compressor blade or aircraft wing; 2. Static pressure chamber; 3. Static pressure gap on the pressure surface; 4. Static pressure gap on the suction surface; 5. Slider; 6. Vortex generator; 7. Spring chamber; 8. Spring. Detailed Implementation

[0021] The present invention will be further described below with reference to the accompanying drawings. The following embodiments are only used to more clearly illustrate the technical solution of the present invention, and should not be used to limit the scope of protection of the present invention.

[0022] Example 1

[0023] like Figure 1 As shown, this embodiment introduces a blade or wing with a built-in adaptive vortex generator. Its structure includes a compressor blade or wing 1, a static pressure chamber 2, a pressure surface static pressure slot 3, a suction surface static pressure slot 4, a slider 5, and a vortex generator 6 (its distribution on the slider is as shown in the diagram). Figure 3As shown), spring cavity 7 and spring 8. The static pressure cavity 2 is a hollow structure within the blade or wing 1. The pressure surface static pressure slot 3 is located on the pressure surface of the blade or wing 1, and the suction surface static pressure slot 4 is located on the suction surface of the blade or wing 1. A movable slider 5 is built into the static pressure cavity 2, and an eddy current generator 6 is mounted on the slider 5. The static pressure cavity 2 is connected to the hollow spring cavity 7, and the slider 5 is connected to the wall of the spring cavity via spring 8. The suction surface static pressure slot 4 and the pressure surface static pressure slot 3 are used to conduct the static pressure on the suction and pressure surfaces of the blade or wing 1 to the upper and lower sides of the slider 5, respectively. When the blade or wing 1 is at a small angle of attack α and there is no suction surface flow separation (i.e.... Figure 2 (As shown in the diagram) The pressure difference between the suction and pressure surfaces of the blade or wing 1 is low, therefore the pressure difference between the upper and lower sides of the slider 5 is also low. At this time, the eddy current generator 6 is completely inside the static pressure chamber 2 and is in a non-working state (e.g.) Figure 5 As shown), this avoids the additional drag and vortex drag generated by the vortex generator 6; and when the blade or wing 1 is at a large angle of attack α and there is suction surface flow separation (i.e. Figure 1 (As shown in the diagram) The pressure difference between the upper and lower sides of slider 5 is relatively high, causing slider 5 to move towards the suction surface, which drives vortex generator 6 to extend out of the suction surface of the blade or wing 1, thus putting vortex generator 6 into working condition (e.g. Figure 4 As shown in the figure, it can effectively suppress suction surface flow separation under high angle of attack conditions and improve the aerodynamic performance of compressors or wings.

[0024] The height h of the vortex generator 6 is 5% of the maximum thickness H of the compressor blade or aircraft wing 1. ～ 50%, the height h of the vortex generator 6 and other geometric parameters are determined by optimization based on the maximum angle of attack of the compressor blade or the aircraft wing 1.

[0025] The height of the static pressure chamber 2, H2 = H1 + h - H3, is where H1 is the height of the slider 5 and H3 is the height of the static pressure slot 4 on the suction surface. Assuming the compressor blade or aircraft wing 1 is at a critical separation angle of attack of αc at a certain incoming flow velocity, the pressure difference between the static pressure slot 3 on the pressure surface and the static pressure slot 4 on the suction surface is ΔP1; while when the compressor blade or aircraft wing 1 is at its maximum operating angle of attack αm, the pressure difference between the static pressure slot on the pressure surface and the static pressure slot on the suction surface is ΔP2. When the compressor blade or aircraft wing 1 is in a non-separated state (i.e., the pressure difference between the static pressure gap 3 on the pressure surface and the static pressure gap 4 on the suction surface ΔP < ΔP1, and the angle of attack α < αc), the lower surface of the slider 5 is in close contact with the lower surface of the static pressure chamber 2. In this non-separated state, the slider 5 does not move, avoiding losses caused by additional mechanical movement. When the compressor blade or aircraft wing 1 is at the maximum working angle of attack (the pressure difference between the static pressure gap 3 on the pressure surface and the static pressure gap 4 on the suction surface ΔP = ΔP2, and the angle of attack α = αm), the vortex generator 6 is in a fully extended state, and at this time, the upper surface of the slider 5 is in close contact with the upper surface of the static pressure chamber 2.

[0026] The preload F0 (direction pointing towards the pressure surface) and elastic coefficient k of the spring 8 are given by design. Specifically, assuming the spanwise thickness of the slider 5 is w, when the eddy current generator 6 is in a fully extended state (the lower surface of the slider 5 is in close contact with the lower surface of the static pressure chamber 2), F0 = bwΔP1, and the elastic coefficient k = bw(ΔP2-ΔP1) / (H2-H1), where b is the width of the slider 5.

[0027] The following description, in conjunction with a preferred embodiment, illustrates the content involved in the above embodiments.

[0028] like Figure 1 As shown, for a typical low-speed compressor stator blade with high diffusion, the chord length is 60 mm, the maximum thickness is 6 mm, and the blade spanwise height is 10 mm. The bleed inlet and jet inlet are designed at approximately 60% of the chord length from the leading edge. When the inlet Mach number is 0.1 and the angle of attack α≈5°, the compressor blades are just about to separate, at which point the static pressure difference coefficient ΔCp≈0.3 between the bleed inlet and jet inlet. However, under the maximum operating angle of attack (α≈15°, with suction surface separation), the static pressure difference coefficient ΔCp≈0.8 between the bleed inlet and jet inlet is selected. The vortex generator 6 height h=2.5 mm, the static pressure chamber 2 height H2=5 mm, the slider 5 height H1=3 mm, the suction surface static pressure slot 4 height H3=0.5 mm, the slider 5 width b=4 mm, and the slider 5 spanwise thickness w=10 mm. It can be estimated that for α≈5°, the pressure difference ΔP between the upper and lower surfaces of slider 5 is ΔP≈208Pa; for α≈15°, the pressure difference ΔP between the upper and lower surfaces of slider 5 is ΔP≈554Pa. According to the relevant formulas and explanations of the claims and specific implementation measures, the preload force F0≈0.08N and the elastic coefficient k≈69N / M of spring 8 can be calculated.

[0029] The working principle of a blade or wing that suppresses flow separation through a built-in vortex generator 6 is as follows: The pressure surface static pressure slot 3 of this device is located at the pressure surface of the compressor blade or aircraft wing 1 where the pressure is higher, while the suction surface static pressure slot 4 is located at the suction surface of the blade or wing. Therefore, there is a certain pressure difference between the pressure surface and the suction surface. When the compressor blade or aircraft wing 1 is at different angles of attack (with the incoming flow velocity remaining constant), the pressure difference between the pressure surface and the suction surface will change accordingly. When the angle of attack is small, the pressure difference between the pressure surface and the suction surface is small, resulting in a low pressure difference on both sides of the slider 5. Due to the preload of the spring 8, the slider 5 is in a closed state against the lower surface of the static pressure chamber 2. The vortex generator 6 does not extend out of the blade or wing surface, thus having no flow control effect and not generating additional type drag or vortex drag. As the angle of attack increases, the pressure difference between the pressure surface and the suction surface gradually increases. After the pressure difference on both sides of the slider 5 cancels the preload generated by the spring 8, it will push the slider 5 and cause the vortex generator 6 to extend out of the blade or wing surface, thereby suppressing the flow separation on the suction surface under high angle of attack conditions and improving the aerodynamic performance of the compressor blade or aircraft wing 1.

[0030] The innovation of this invention lies in its ingenious solution to the problem that while flow control is required to activate in compressor blades or aircraft wings 1 under flow separation conditions, activating the flow controller leads to greater losses under non-flow separation conditions. The pressure difference between the pressure and suction surfaces of the compressor blades or aircraft wings 1 can reflect different operating conditions (different angles of attack) and also serve as the power source for the release and recovery of the vortex generator 6. Based on these principles, this invention can achieve adaptive start-stop of the vortex generator 6 on the compressor blades or aircraft wings 1 with a relatively simple structure, without requiring an external air source, power supply, or related auxiliary systems.

[0031] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the technical principles of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A blade or wing with a built-in adaptive eddy current generator, characterized in that, include: The compressor blade or aircraft wing (1), static pressure chamber (2), pressure surface static pressure slot (3), suction surface static pressure slot (4), slider (5), vortex generator (6), spring chamber (7), and spring (8); wherein, the static pressure chamber (2) is a hollow structure inside the compressor blade or aircraft wing (1), the pressure surface static pressure slot (3) is located on the pressure surface of the compressor blade or aircraft wing (1), the suction surface static pressure slot (4) is located on the suction surface of the compressor blade or aircraft wing (1), the static pressure chamber (2) contains a movable slider (5), the slider (5) is equipped with a vortex generator (6) on the side facing the suction surface static pressure slot (4), the vortex generator (6) can extend through the suction surface static pressure slot (4) to the outside of the compressor blade or aircraft wing (1), the static pressure chamber (2) and the hollow spring chamber (8) 7) Connected, the slider (5) is connected to the wall of the spring cavity through the spring (8). The static pressure seam (3) and the static pressure seam (4) of the pressure surface are respectively used to conduct the static pressure of the suction surface and the pressure surface of the compressor blade or wing (1) to the upper and lower surfaces of the slider (5). When the compressor blade or wing (1) is in a small angle of attack and there is no suction surface flow separation, the pressure difference between the upper and lower sides of the slider (5) is low. At this time, the vortex generator (6) is completely inside the static pressure cavity (2) and is in a non-working state. When the compressor blade or wing (1) is in a large angle of attack and there is suction surface flow separation, the pressure difference between the upper and lower sides of the slider (5) is high, which makes the slider (5) move towards the suction surface and drive the vortex generator (6) to extend out of the suction surface of the compressor blade or wing (1), so that the vortex generator (6) is in a working state.

2. The blade or wing with a built-in adaptive eddy current generator according to claim 1, characterized in that, The height h of the vortex generator (6) is 5% of the maximum thickness H of the compressor blade or aircraft wing (1). ～ 50%.

3. The blade or wing with a built-in adaptive eddy current generator according to claim 1, characterized in that, The height of the static pressure chamber (2) is H2 = H1 + h - H3, where H1 is the height of the slider (5) and H3 is the height of the static pressure gap (4) on the suction surface.

4. The blade or wing with a built-in adaptive eddy current generator according to claim 3, characterized in that, Assuming that the compressor blades or aircraft wings (1) are in the critical state of just separation and the maximum working angle of attack, the pressure difference between the static pressure gap (3) on the pressure surface and the static pressure gap (4) on the suction surface is △P1 and △P2 respectively, △P1 < △P2; when the actual pressure difference between the static pressure gap (3) on the pressure surface and the static pressure gap (4) on the suction surface is △P≤△P1, the lower surface of the slider (5) is close to the lower surface of the static pressure chamber (2), and the vortex generator (6) is in the unextended state; when △P1 < △P < △P2, the lower surface of the slider (5) leaves the lower surface of the static pressure chamber (2), and the vortex generator (6) is in the partially extended state; when △P = △P2, the upper surface of the slider (5) is close to the upper surface of the static pressure chamber (2), and the vortex generator (6) is in the fully extended state.

5. The blade or wing with a built-in adaptive eddy current generator according to claim 4, characterized in that, Assuming the spanwise thickness of the slider (5) is w, when the eddy current generator (6) is in a fully extended state, the preload force F0 of the spring (8) is calculated as follows: F0 = bwΔP1, where b is the width of the slider (5).

6. The blade or wing with a built-in adaptive eddy current generator according to claim 5, characterized in that, The formula for calculating the elastic coefficient k of the spring (8) is as follows: k = bw(△P2-△P1) / (H2-H1).

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