A cross-speed regime variant aircraft

Abstract

The application discloses a transonic variable aircraft, and relates to the field of aircrafts; the aircraft comprises a fuselage and a wing in a blended wing body design structure, the wing comprises an outer wing and an inner wing, and the sweepback angle of the outer wing and the folding angle of the inner wing can be adjusted, so that the aircraft can present a low-speed configuration, a high-speed configuration and multiple intermediate configurations between the low-speed configuration and the high-speed configuration to adapt to multiple inflow conditions such as transonic speed, supersonic speed and the like. The transonic variable aircraft design method provided by the application can adapt to different speed domains by changing the aerodynamic shape of the aircraft.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a cross-speed-range variant aircraft. Background Technology

[0002] The development of modern aircraft originated from biomimetic research on bird flight patterns. Based on the wing attitudes and cross-sectional shapes of birds during gliding, various aircraft adapted to different flight states have been designed. However, whether fixed-wing or rotary-wing, the aerodynamic configuration of an aircraft does not change significantly during flight. Therefore, an aircraft typically has only one cruise mode and cannot achieve efficient flight across speed ranges. Birds, on the other hand, can adjust their shape and attitude to achieve optimal flight performance according to different flight states, achieving a perfect balance between efficiency and performance. Therefore, based on biomimetic research on bird variant flight, designing and studying cross-speed-range variant aircraft that can adapt to different speed ranges by changing their aerodynamic shape provides design ideas and theoretical foundations for the future development of aircraft, and has significant academic and engineering value. Summary of the Invention

[0003] The purpose of this invention is to provide a cross-speed-range variant aircraft to solve the problems existing in the prior art. By changing the aerodynamic shape of the aircraft itself, it can adapt to flight in different speed ranges.

[0004] To achieve the above objectives, the present invention provides the following solution:

[0005] This invention provides a transvestite aircraft, comprising a fuselage and a wing with a blended wing-body structure. The wing includes an outer wing and an inner wing, and the sweep angle of the outer wing and the folding angle of the inner wing are both adjustable. This allows the aircraft to present low-speed configuration, high-speed configuration, and various intermediate configurations between low-speed and high-speed configurations to adapt to various incoming flow conditions such as low speed, subsonic, transonic, and supersonic.

[0006] Optionally, when the wings are fully extended, the second sweep angle of the outer wing can be at its minimum, and the wing's projected area and aspect ratio can both reach their maximum, thus enabling the aircraft to adapt to low-speed flight.

[0007] Optionally, during wing deformation, the inner wing, under the action of the deformation mechanism, can fold upward around a pivot parallel to the fuselage axis to increase the dihedral angle of the wing, thereby reducing the wing's projected area and aspect ratio. The first section of the outer wing can drive the second section of the outer wing to fold downward to maintain lift. At the same time, the sweep angle of the second section of the outer wing is increased by the variable sweep angle device between the first and second sections of the outer wing, thereby improving the high-speed performance of the aircraft. After the wing deformation is completed, the inner wing is completely attached to the fuselage sidewall, and the overall wing projected area and aspect ratio of the aircraft reach their minimum.

[0008] In typical aerodynamic configurations, the flying wing design has been used on long-range bombers with stealth penetration capabilities and some unmanned combat aerial vehicles (UCAVs), allowing them to appear undetected behind enemy lines and deliver the most lethal blows. A significant advantage of the flying wing design is its higher aerodynamic lift efficiency compared to other configurations. This stems from two aspects: first, in a flying wing, the wing and fuselage are largely integrated, with the fuselage contributing to lift, whereas conventional fuselages contribute very little; second, the flying wing design uses only a single wing surface, avoiding lift loss caused by unfavorable interference between wing surfaces in conventional configurations. Therefore, the wing adopts a flying wing configuration, with a supersonic double-arc symmetrical airfoil. The leading edge sweep angle varies from 10 to 50°, the aspect ratio from 3.3 to 4.9, and the wing reference area from 49.5 to 145.7 m². 2 The wing loading range based on the maximum takeoff weight is 130.4–383.3 kg / m. 2 ;

[0009] When an aircraft is in steady level flight, it moves at a constant velocity in a straight line, as shown by the following formula.

[0010]

[0011]

[0012] The level flight speed of the aircraft is obtained as

[0013]

[0014] Where L is lift; G is the aircraft's weight; T is the engine thrust; D is drag; and S is the thrust. ref —Aircraft reference area, C L —Lift coefficient; C D —Drag coefficient; To maintain the aircraft configuration with minimum drag, the level flight speed V also reaches its maximum when the aircraft engine thrust reaches its maximum, i.e., the maximum level flight speed V. max Maintaining the aircraft configuration to maximize lift will minimize the aircraft's level flight speed, i.e., the minimum level flight speed V. min At different altitudes, the maximum level flight speed V of the aircraft max and minimum level flight speed V min The possible level flight speed range between these parameters becomes the speed envelope of the aircraft. Based on the above, when the takeoff weight and engine of the aircraft are selected, the level flight performance of the aircraft is determined by its aerodynamic coefficient and reference area, which is the aerodynamic shape of the aircraft.

[0015] For different flight Mach numbers, a small sweep angle and a large aspect ratio are used during takeoff, landing, and low-speed flight; as the flight speed increases, the sweep angle of the aircraft is increased and the aspect ratio is decreased; when the flight speed reaches subsonic speed, the sweep angle is further increased; when the flight speed reaches supersonic speed, a fuselage with a large aspect ratio and a delta wing are adopted.

[0016] When the aircraft is flying at low speed, the wings are fully extended, and the sweep angle of the second section of the outer wing is at its minimum. At this time, the wing's projected area and aspect ratio are at their maximum. When the flight speed increases, the inner wing, under the action of the deformation mechanism, folds upward around a pivot parallel to the fuselage axis to increase the wing's dihedral angle and decrease the wing's projected area and aspect ratio. The first section of the outer wing drives the second section of the outer wing to fold downward to maintain lift. At the same time, the sweep angle of the second section of the outer wing is increased by the variable sweep angle device between the first and second sections of the outer wing, improving the aircraft's high-speed performance. When the deformation is complete, the inner wing is completely attached to the fuselage sidewall and no longer provides lift. The wing area and aspect ratio of the entire aircraft reach their minimum, and the aircraft enters the high-speed deformation state.

[0017] The relative thickness distribution of the outer wing is as follows The wing's maximum thickness is 3.375%, located at the 50% chord length. A major drawback of this airfoil is its sharp leading edge, making it prone to separation and resulting in poor low-speed performance. Since the aircraft must use this airfoil for low-speed flight, it will cause some performance loss. In this paper, to balance the aircraft's low- and high-speed performance, some compromises must be made. This airfoil is selected as the initial option and can be adjusted based on later calculations. Furthermore, with the deepening research on adaptive flexible deformable wings, this technology could be considered for future applications in morphing aircraft, avoiding similar compromises and further improving the aircraft's adaptability.

[0018] Regarding the selection of the inner wing root airfoil, considering that the inner wing adheres to the fuselage and does not provide lift during supersonic flight, and that the actuators and pivots required for wing deformation need to be housed within it, its relative thickness does not need to be too small. However, to reduce the interference of its folding on the surrounding flow field, an airfoil with excessively large relative thickness is also not advisable. In this invention, the widely used NACA63A006 airfoil is adopted for the inner wing root.

[0019] The present invention achieves the following technical effects compared to the prior art:

[0020] Based on the requirements of the transvestite (Mach number 0.3-3.0) flight performance indicators for novel variant aircraft, this invention proposes a novel aerodynamic layout and variant design scheme from the perspective of aircraft aerodynamic layout design. The design scheme includes a low-speed configuration adapted to an incoming flow Mach number of 0.3 and a high-speed configuration adapted to an incoming flow Mach number of 3. Furthermore, by further adjusting the deformation parameters such as the outer wing sweep angle and inner wing folding angle, the aircraft can present multiple intermediate configurations between the low-speed and high-speed configurations to adapt to various incoming flow conditions such as transonic and supersonic speeds. Based on this scheme, computational fluid dynamics is used to calculate the longitudinal aerodynamic characteristics of different variant configurations of the aircraft within the transvestite flight Mach number range. Attached Figure Description

[0021] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0022] Figure 1 This is a schematic diagram of the planar shape of the wing of the present invention when it is retracted;

[0023] Figure 2 This is a schematic diagram of the planar shape of the wing of the present invention when it is deployed;

[0024] Figure 3 A schematic diagram showing the distribution of airfoil data for the wing roots of the outer and inner wings;

[0025] Figure 4 This is a schematic diagram of cross-sections at seven different locations along the fuselage axis of the aircraft.

[0026] Figure 5a This is a graph showing the lift coefficient in the aerodynamic characteristic curves of an aircraft in a low-speed configuration.

[0027] Figure 5b This is a plot of the drag coefficient in the aerodynamic characteristic curves of an aircraft in a low-speed configuration.

[0028] Figure 5c This is a schematic diagram of the lift-to-drag ratio in the aerodynamic characteristic curve of an aircraft in a low-speed configuration.

[0029] Figure 6a This is a graph showing the lift coefficient in the aerodynamic characteristic curves of an aircraft in a high-speed configuration.

[0030] Figure 6b This is a plot of the drag coefficient in the aerodynamic characteristic curves of an aircraft in a high-speed configuration.

[0031] Figure 6c This is a schematic diagram of the lift-to-drag ratio in the aerodynamic characteristic curve of an aircraft in a high-speed configuration.

[0032] Figure 7a The image shows the lift coefficient in the longitudinal aerodynamic characteristic curve of the aircraft at transonic speed Ma = 0.85.

[0033] Figure 7b The plot shows the drag coefficient in the longitudinal aerodynamic characteristic curve of the aircraft at transonic speed Ma = 0.85.

[0034] Figure 7c This is a schematic diagram of the lift-to-drag ratio in the longitudinal aerodynamic characteristic curve of an aircraft at transonic speed Ma = 0.85.

[0035] Figure 8a The lift coefficient diagram is shown in the longitudinal aerodynamic characteristic curve of the aircraft at supersonic speed Ma=1.5.

[0036] Figure 8b The image shows the drag coefficient in the longitudinal aerodynamic characteristic curve of the aircraft at supersonic speed Ma = 1.5.

[0037] Figure 8c This is a schematic diagram of the lift-to-drag ratio in the longitudinal aerodynamic characteristic curve of an aircraft at supersonic speed Ma=1.5.

[0038] Explanation of reference numerals in the attached diagram: 1-fuselage, 2-inner wing, 3-outer wing. Detailed Implementation

[0039] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0040] The purpose of this invention is to provide a cross-speed-range variant aircraft to solve the problems existing in the prior art. By changing the aerodynamic shape of the aircraft itself, it can adapt to flight in different speed ranges.

[0041] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0042] This invention provides a cross-velocity-range variant aircraft, such as Figure 1 and Figure 2As shown, the aircraft includes a fuselage 1 and a wing employing a blended wing-body structure. The wing comprises an inner wing 2 and an outer wing 3, with the outer wing 3 consisting of a first section and a second section. The sweep angle of the inner wing 2 and the folding angle of the outer wing 3 are both adjustable, allowing the aircraft to adopt low-speed configurations, high-speed configurations, and various intermediate configurations between low-speed and high-speed configurations to adapt to various incoming airflow conditions, including low-speed, subsonic, transonic, and supersonic. When the wing is fully extended, the sweep angle of the second section of the outer wing 3 is at its minimum, and the wing's projected area and aspect ratio are at their maximum, enabling the aircraft to adapt to low-speed flight. During wing deformation, the inner wing 2, under the action of the deformation mechanism, can fold upward around a pivot parallel to the fuselage axis to increase the dihedral angle of the wing, reduce the wing's projected area and aspect ratio. The first section of the outer wing 3 can drive the second section of the outer wing 3 to fold downward to maintain lift. At the same time, the sweep angle of the second section of the outer wing 3 is increased by the variable sweep angle device between the first and second sections of the outer wing 3, improving the high-speed performance of the aircraft. After the wing deformation is completed, the inner wing 2 is completely attached to the fuselage sidewall, and the total wing projected area and aspect ratio of the aircraft reach their minimum.

[0043] This invention provides a design process for a cross-speed-range variability aircraft: First, the design requirements of the aircraft are clearly defined. Based on the aircraft's aerodynamic coefficient and reference area, the level flight performance requirements are determined. According to performance indicators, the variability aircraft must possess good aerodynamic performance across different level flight Mach number ranges (0.3-3.0), enabling it to switch between multiple mission roles, such as low-speed, long-endurance cruising and cruise, and high-speed strike and penetration. According to classical flight mechanics theory, an aircraft flies in a straight line at a constant speed during cruise. The level flight performance of an aircraft is typically determined by its maximum level flight speed V. max and minimum level flight speed V minThe evaluation is based on the possible level flight speed range; the design of aircraft variant parameters is performed, and variant aircraft adapt to different flight environments and improve aerodynamic characteristics by changing their shape. Generally speaking, wing shape can be varied at three scales: large-scale deformation, which is the change in the area, sweep angle, and cross-sectional shape of the entire wing; small-scale deformation, which is the change in the shape of local parts of the wing, such as local bulges, variable leading edges, adaptive wingtips, etc. Medium-scale and small-scale deformations can mostly only improve the aerodynamic efficiency of the wing at a single design point, while large-scale deformation can achieve significant benefits for different flight conditions. Based on this, we can select the main aircraft variant parameters of interest as the master control parameters for the design of variant schemes; for different flight Mach numbers, the direction of aircraft deformation is determined to change the aspect ratio and outer wing sweep angle; the initial parameters of the aircraft are determined, including the range of parameters for takeoff weight, empty weight, fuel coefficient, takeoff thrust-to-weight ratio, and wing loading; the aerodynamic layout and wing design are determined, and when selecting the aerodynamic layout of the aircraft, we mainly consider the main purpose of the aircraft. The variant aircraft designed in this paper mainly emphasizes its long-endurance patrol and standby capabilities and high-speed penetration performance. Long endurance requires the aircraft to have a high lift-to-drag ratio, which requires the largest possible maximum lift coefficient; high-speed penetration requires the aircraft to have good high-speed and stealth performance.

[0044] In typical aerodynamic configurations, the flying wing design has been used on long-range bombers with stealth penetration capabilities and some unmanned combat aerial vehicles (UCAVs). These aircraft can appear unnoticed behind enemy lines and deliver the most lethal blows. A significant advantage of the flying wing design is its higher aerodynamic lift efficiency compared to other configurations. This stems from two aspects: first, in a flying wing design, the wing and fuselage are essentially integrated, with the fuselage contributing to lift, whereas conventional fuselages generate very little lift; second, the flying wing design uses only a single wing surface, avoiding lift loss caused by unfavorable interference between wing surfaces in conventional configurations. Therefore, the wing adopts a flying wing configuration, with the airfoil using a supersonic double-arc symmetrical airfoil.

[0045] Specifically, the takeoff weight of the aircraft is the first parameter that needs to be determined in the overall design. This embodiment is positioned as a single-engine light unmanned aerial vehicle. Referring to existing designs and taking into account the payload, fuel, and empty weight, the maximum takeoff weight is initially estimated to be 19t, the empty weight is 8t, and the fuel coefficient is tentatively set at 0.4.

[0046] To meet the engine thrust requirements during high Mach number cruise, the aircraft's takeoff thrust-to-weight ratio is designed to be 0.8, which translates to a maximum afterburner thrust in the range of 150 kN. A certain turbofan engine achieves this range, with the following main parameters: length 4.826 m, maximum diameter 1.13 m, weight 1360 kg, maximum thrust 100 kN, afterburner thrust 156 kN, and a thrust-to-weight ratio of 11.7. This embodiment is based on the above parameters for subsequent design.

[0047] Wing loading is also a key parameter determining aircraft performance and maneuverability. In this design, the aircraft shape will change for different flight missions, meaning the wing reference area is not a fixed value, resulting in varying wing loading. Therefore, in determining the initial parameters, we can only provide a range for wing loading based on existing aircraft parameters: 120-400 kg / m². 2 .

[0048] The aircraft has a maximum takeoff weight of 19 tons and a wing loading of 150-400 kg / m. 2 Using these as basic parameters, we designed the wing. When the wing is fully extended, the aircraft becomes a wing-body flying wing configuration, with the fuselage also providing lift. Therefore, when calculating the wing reference area, we also considered the projected area of ​​the fuselage. However, when the wing is folded upwards and retracted, the aircraft can no longer be strictly considered a flying wing configuration; therefore, we did not consider the fuselage when calculating the wing reference area. The wing leading edge sweep angle varies from 10 to 50°, the aspect ratio varies from 3.3 to 4.9, and the wing reference area varies from 49.5 to 145.7 m. 2 The wing loading varies from 130.4 to 383.3 kg / m based on the maximum takeoff weight. 2 .

[0049] Regarding the selection of the inner wing root airfoil, considering that the inner wing will not provide lift during supersonic flight as it will be flush with the fuselage, and that the actuators and pivots required for wing deformation need to be housed within it, its relative thickness does not need to be too small. However, to reduce the interference of its folding on the surrounding flow field, an airfoil with excessively large relative thickness is also not advisable. In this design, the widely used NACA63A006 airfoil is adopted for the inner wing root. The airfoil data for the outer and inner wing roots are as follows: Figure 3 As shown.

[0050] When designing the fuselage, we selected seven cross-sections at different locations along the fuselage axis for parametric design, such as... Figure 4 As shown, the fuselage is integrally formed by sweeping multiple sections, ensuring its smoothness. The design dimensions of each section are shown in Table 1.

[0051] Table 1 Design Dimensions of Each Section

[0052]

[0053] The main aerodynamic layout parameters of the aircraft when the wings are deployed and retracted are as follows:

[0054] (1) When the wings are fully extended and the sweep angle is minimal:

[0055] Theoretical root chord length (fuselage length): 15.151m

[0056] Wingtip chord length (excluding wingtip): 1.48m

[0057] Root-to-shoot ratio: 10.24

[0058] Wing area: 145.7m² 2

[0059] Wetting area: 307.01m 2

[0060] Wingspan: 26.65m

[0061] Aspect ratio: 4.9

[0062] Nose sweep angle: 55°

[0063] Leading edge sweep angle of the second wing segment: 10°

[0064] Fuselage length: 15.5m

[0065] (2) When the wings are fully retracted and the sweep angle is at its maximum:

[0066] Theoretical root chord length: 7.866m

[0067] Wingtip chord length (excluding wingtip): 1.8m

[0068] Root-to-shoot ratio: 4.37

[0069] Wing area: 49.5m² 2

[0070] Wetting area: 236.668m² 2

[0071] Wingspan: 13.05m

[0072] Aspect ratio: 3.3

[0073] Nose sweep angle: 55°

[0074] Wing sweep angle at 40% chord position: 43°

[0075] Fuselage length: 15.5m

[0076] Numerical calculation of aircraft aerodynamic characteristics

[0077] To verify whether the initial layout scheme of the above-mentioned aircraft meets the aerodynamic characteristics required for cross-speed range flight, the aerodynamic parameters of the two configurations of the variant aircraft at their corresponding Mach numbers were calculated.

[0078] (1) Low-speed configuration

[0079] The variant aircraft's configuration during cruise at Ma=0.3 is with its folding wings fully extended and its sweep angle at its minimum.

[0080] Figure 5a , Figure 5b and Figure 5c The figure shows the aerodynamic characteristics of the variant aircraft in a low-speed configuration (i.e., fully deployed wing configuration) calculated numerically. From 5a, Figure 5b It can be seen that within an angle of attack of 10°, the lift coefficient of the aircraft increases approximately linearly with the increase of the angle of attack, while the drag coefficient increases slowly. However, when the angle of attack exceeds 10°, the slope of the lift line gradually decreases, and the growth slows down, while the drag growth rate increases significantly. This is due to the negative impact of using a sharp leading-edge airfoil, because at larger angles of attack, flow separation is more likely to occur at the leading edge of the airfoil. Figure 5c The aircraft's lift-to-drag ratio reaches its maximum near a 2-degree angle of attack, with a maximum lift-to-drag ratio of approximately 15, after which it decreases as the angle of attack increases. Furthermore, based on a maximum design takeoff weight of 19 tons, the lift coefficient required for level flight at an altitude of 3 km and a cruise speed of Ma = 0.3 is 0.299. Figure 5a The variant aircraft achieved lift coefficients of 0.24933 and 0.352 at angles of attack of 4 and 6 degrees, respectively, indicating that its level flight angle of attack is between 4 and 6 degrees. Furthermore, the aircraft's drag coefficient at an angle of attack of 6 degrees is 0.035, equivalent to an aerodynamic drag of 22.6 kN, demonstrating that the selected engine fully meets the thrust requirements of the aircraft. Additionally, the lift-to-drag ratio at an angle of attack of 6 degrees is 10.1, indicating good cruise economy.

[0081] (2) High-speed configuration

[0082] As required, the morphing aircraft should have cruise performance at Ma=3.0 at high speeds. Due to the generation of shock waves, the aircraft will experience significant shock wave drag during supersonic flight. To reduce wave drag, the aircraft should adopt a configuration with a small aspect ratio and a large sweep angle. Based on this, we define the morphing aircraft's configuration at Ma=3.0 as the morphing state when the folding wings are fully retracted and the sweep angle reaches its maximum position.

[0083] Figure 6a , Figure 6b and Figure 6c The aerodynamic characteristics of the variant aircraft in high-speed configuration are shown through numerical calculations. Figure 6a It can be seen that at a 0° angle of attack, the lift coefficient is negative, and thereafter, as the angle of attack increases, the lift coefficient of the aircraft increases approximately linearly. From Figure 6bIt can be seen that under small angles of attack, the drag coefficient of the aircraft does not change much with the angle of attack. However, after the angle of attack exceeds 8 degrees, the drag coefficient increases rapidly with the angle of attack. This is mainly because the aircraft's frontal area increases rapidly with the angle of attack, and the shock wave causes the flow to decelerate and pressurize on the aircraft's frontal side, resulting in a larger pressure difference between the frontal and leeward sides. The aircraft's lift-to-drag ratio reaches its maximum value in the range of 10-15 degrees of angle of attack, with a maximum lift-to-drag ratio of approximately 2.8. If calculated based on a maximum design takeoff weight of 19t, the required lift coefficient at an altitude of 21km and Ma=3.0 is 0.126. Figure 5a The lift coefficients of the vari-engine aircraft at angles of attack of 4 degrees and 6 degrees are 0.079 and 0.142, respectively, indicating that the angle of attack for level flight is between 4° and 6°. The drag coefficient at an angle of attack of 6 degrees is 0.073, equivalent to an aerodynamic drag of 108 kN, which is 70% of the maximum afterburner thrust. This demonstrates that the selected engine meets the thrust requirements for level flight. Furthermore, the above calculations are based on the aircraft under maximum load conditions. In reality, aircraft flying at maximum speed are often engaged in attack or escape missions, and the load is certainly less than the maximum weight, indicating a lower lift requirement. Therefore, as long as the selected engine meets the requirements, the vari-engine aircraft can achieve high-speed flight at higher altitudes.

[0084] (3) Intermediate configuration

[0085] Based on the aerodynamic characteristics analysis of low-speed and high-speed configurations above, this invention calculates and analyzes the aerodynamic characteristics of different intermediate configurations formed by combining two deformation parameters, namely fold angle and sweep angle, under two incoming flow Mach numbers of Ma = 0.85 and 1.5. Nine typical state examples are provided in this invention, defined as follows: State U has a fold angle of 0°, including: U1 state: sweep angle 10°, U2 state: sweep angle 30°; State M has a fold angle of 45°, including: M1 state: sweep angle 10°, M2 state: sweep angle 30°, M3 state: sweep angle 50°; State F has a fold angle of 135°, including: F1 state: sweep angle 10°, F2 state: sweep angle 30°, F3 state: sweep angle 50°.

[0086] Under transonic flight conditions (Ma = 0.85), the equivalent aerodynamic characteristics of the morphing vehicle in different deformation modes are as follows: Figure 7a , Figure 7b and Figure 7c As shown, the aerodynamic coefficients for each operating condition are uniformly adopted using the projected area of ​​the wing plane in the U1 state as a dimensionless reference value. Figure 7a , Figure 7b and Figure 7cIt can be seen that there are significant differences in the aerodynamic performance of the aircraft under different deformation modes, indicating that the variant scheme proposed in this paper can meet the requirements of changing the aerodynamic performance of the aircraft as needed.

[0087] The equivalent lift coefficient of the aircraft varies with the angle of attack under different deformation states, as follows: Figure 7a As shown, overall, the deformation parameter that has the greatest impact on the lift characteristics of an aircraft is the wing folding angle. The lift coefficients at different sweep angles under the three folding angle states (U, M, and F) are not significantly different, forming three distinct families of curves. At 0 degrees of angle of attack, the lift coefficients of different sweep angle configurations are almost identical under the U and M folding angle states. This is because the flow is parallel to the fuselage axis at this point, and the wing sections are symmetrical airfoils that generate almost no lift. Most of the lift is due to the flow differences caused by the different geometries of the upper and lower surfaces of the fuselage. Similarly, the same lift coefficient also appears among the three different sweep angles when the wing is fully folded (F state). Because the aerodynamic performance under supersonic flight conditions was considered during the design, the second section of the outer wing of the aircraft is selected as a double-arc airfoil with a very sharp leading edge and a very thin thickness. This leads to easy separation of the flow at the leading edge of the wing, which is also reflected in the lift characteristics of the aircraft. Due to the nonlinear change in lift coefficient caused by flow separation, the lift coefficients in states 1 and 2 with smaller sweep angles cease linear growth at a 4-degree angle of attack, while the lift in state 3 with a larger sweep angle also ceases linear growth after an 8-degree angle of attack. It is noteworthy that in states U1 and M1, the increase in lift coefficient with increasing angle of attack slows significantly after 4 degrees, causing it to be surpassed by states U2 and M2. Therefore, after an angle of attack greater than 6 degrees, U2 and M2 become the conditions with the highest lift coefficients for their respective folding angles. For state F with the wing fully folded, the lift coefficient at F1 is slightly greater than that at all angles of attack than the other two sweep angle cases, F2 and F3.

[0088] The equivalent drag coefficient of an aircraft varies with angle of attack under different deformation states, as follows: Figure 7b As shown, similar to the lift characteristics, within the angle of attack range of less than 4 degrees, the difference in equivalent drag coefficient between different deformation states is very small, and the drag values ​​are all less than 0.05, demonstrating the excellent low-drag characteristics of the designed aircraft during transonic flight. Beyond 4 degrees of angle of attack, the equivalent drag coefficient of the aircraft increases significantly with increasing angle of attack, and the drag increment between different folding angle states is almost in line with the size of their corresponding actual projected area, i.e., states U, M, and F in sequence. Furthermore, we note that the state with the largest drag among all folding angle states is state 1 with the smallest sweep angle, followed by states 2 and 3. This is consistent with the theory in aircraft aerodynamics that as the wing sweep angle increases during transonic flight, drag decreases.

[0089] The lift-to-drag ratio of an aircraft varies with the angle of attack as follows: Figure 7c As shown, at small angles of attack, the three states U1, M1, and F1 with small sweep angles have a larger linear lift line slope than other configurations with the same folding angle. Since the drag coefficients between these states are not significantly different, these three states have a larger lift-to-drag ratio at their respective folding angles. In state U1, at a 2-degree angle of attack, the aircraft's maximum lift-to-drag ratio at Ma = 0.85 can reach 13.75. At larger angles of attack, the increase in lift slows down, while drag increases significantly. At this point, U3, M3, and F3, with the largest sweep angle and the smallest drag coefficient, become the states with the largest lift-to-drag ratio at their respective folding angles, demonstrating the good lift-to-drag characteristics of the large sweep angle configuration at large angles of attack.

[0090] Under supersonic (Ma=1.5) flight conditions, the equivalent aerodynamic characteristics of the morphing vehicle in different deformation modes are as follows: Figure 8a , Figure 8b , Figure 8c As shown in the figure, similar to the case of Ma = 0.85, the differences in aerodynamic performance between different deformation modes, especially at different folding angles, are still quite significant. This indicates that the variant scheme proposed in this paper can also meet the requirements of changing the aerodynamic performance of the aircraft as needed at supersonic speeds.

[0091] The equivalent lift coefficient of the aircraft varies with the angle of attack under different deformation states, as follows: Figure 8a As shown, the lift coefficient curves of the aircraft under nine different deformation states exhibit nine different linear segment slopes, indicating that the lift characteristics of the aircraft during supersonic flight are mainly determined by the deformation folding angle and the outer wing sweep angle. Examining different wing folding states with the same outer wing sweep angle reveals that at the same angle of attack (greater than 0 degrees), the larger the wing folding angle, the smaller the equivalent lift coefficient and the smaller the lift line slope. This is mainly because wing folding primarily affects the actual projected area of ​​the wing; the larger the folding angle, the smaller the projected wing area, thus reducing the aerodynamic loads on the entire aircraft and the increase in aerodynamic loads in the linear segment due to the increased angle of attack. However, when the wing folding state is fixed and different outer wing sweep angles are examined, it can be observed that as the sweep angle increases...

[0092] The lift coefficient and lift slope of the aircraft both decrease, due to the Mach angle of the incoming flow at this time.

[0093]

[0094] The wing leading edge sweep angle varies from 10 to 50 degrees. Therefore, the changes in the aircraft's lift coefficient and slope are likely due to the different supersonic / subsonic flow regimes at different sweep angles, requiring analysis in conjunction with the specific flow structure. Figure 8aIt can also be noted that negative lift occurred in all nine states at a 0-degree angle of attack. This is mainly because during supersonic flight, the airflow is obstructed at the nose, forming a shock wave. The upper surface of the aircraft nose has a larger windward area than the lower surface, resulting in greater airflow obstruction, stronger shock waves, and greater pressure behind the waves. This causes the pressure on the upper surface of the aircraft nose to be greater than that on the lower surface. When a symmetrical wing generates almost no lift at a 0-degree angle of attack, the negative pressure difference between the upper and lower surfaces of the aircraft results in a negative lift coefficient.

[0095] The equivalent drag coefficient of an aircraft varies with angle of attack under different deformation states, as follows: Figure 8b As shown, similar to the transonic speed Ma = 0.85, within the angle of attack range of less than 4 degrees, the difference in equivalent drag coefficient between different deformation states is very small, and the drag values ​​are all less than 0.05, demonstrating that the designed aircraft also possesses good low-drag characteristics during supersonic flight. Beyond 4 degrees of angle of attack, the equivalent drag coefficient of the aircraft increases significantly with increasing angle of attack, and the drag increments between different folding angle states are arranged in descending order of their corresponding actual projected areas U, M, and F states. However, within the same folding angle state, the equivalent drag coefficient decreases with increasing sweep angle, which is due to the fact that a large sweep angle wing can reduce shock wave drag during supersonic flight.

[0096] The lift-to-drag ratio of an aircraft varies with the angle of attack as follows: Figure 8c As shown, comparing different sweep angle states under the same folding angle, although the aerodynamic lift and drag of the aircraft vary significantly with the change in sweep angle, when both are considered together, the differences between the different sweep angle states under the same folding angle are not significant. This indicates that the folding angle is the most important factor affecting the lift-to-drag ratio characteristics of the aircraft at this flight speed. Furthermore, in the folding angles U and M, the aircraft reaches its maximum lift-to-drag ratio near a 6-degree angle of attack, corresponding to values ​​of 4.92 (U1 state) and 4.49 (M1 state), respectively. However, in the folding angle F, the aircraft needs to reach its maximum lift-to-drag ratio near a 10-degree angle of attack, corresponding to a value of 3.16 (F1 state).

[0097] In the description of this invention, it should be noted that the terms "center," "top," "bottom," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0098] Specific examples have been used to illustrate the principles and implementation methods of this invention. The descriptions of the above embodiments are only for the purpose of helping to understand the method and core ideas of this invention. Furthermore, those skilled in the art will recognize that, based on the ideas of this invention, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of this invention.

Claims

1. A transonic variable geometry aircraft, characterized by: The fuselage and wings of the wing-body configuration, the wings including outer wings and inner wings, the sweepback angle of the outer wings and the folding angle of the inner wings being adjustable, so that the aircraft can assume low-speed configuration, high-speed configuration and various intermediate configurations between the low-speed configuration and the high-speed configuration to adapt to low-speed, subsonic, transonic, supersonic and various incoming flow conditions; after the wings are fully deployed, the second section of the outer wings can have a minimum sweepback angle, and the wing projection area and aspect ratio reach a maximum, so that the aircraft can adapt to low-speed flight; During the deformation of the wings, the inner wings can be folded upward around the rotation shaft parallel to the fuselage axis under the action of the deformation mechanism to increase the wing up angle, reduce the wing projection area and aspect ratio, the first section of the outer wings can drive the second section of the outer wings to fold downward to maintain the lift, and at the same time, the sweepback angle between the first section and the second section of the outer wings is changed to increase the sweepback angle of the second section of the outer wings to improve the high-speed performance of the aircraft; after the deformation of the wings is completed, the inner wings are fully attached to the side wall of the fuselage, and the wing projection area and aspect ratio of the whole aircraft reach a minimum; the sweepback angle of the wings ranges from 10° to 50°, and the aspect ratio ranges from 3.3 to 4.9; the maximum thickness of the wings is 3.375%, which is located at 50% of the chord length.

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