Supersonic aircraft

Abstract

A supersonic aircraft capable of achieving low acoustic explosion. The supersonic aircraft comprises a fuselage (1), canards (2), wings (3), empennages (4) and engine nacelles (5). The fuselage (1) is formed in a curved shape in such a manner that a curvature line of the fuselage (1) is in a range between a first end point and a midpoint of a chord line of the fuselage (1) and passes through the chord line at least once from above to below in a lift direction of the supersonic aircraft. The camber line is the center line of the upper contour and the lower contour of the supersonic aircraft in the lift direction when observed from the extension direction of the wings (3) of the supersonic aircraft. The chord line is a line segment formed by connecting the foremost end and the rearmost end of the supersonic aircraft in the flight direction, wherein the foremost end and the rearmost end serve as first end points. Furthermore, when the length of the string is set to be L, the bending line passes through the string at least once from top to bottom in a range from 0.1 L to 0.3 L from the first end point to the middle point in the length direction of the string.

Description

Technical Field

[0001] This application relates to the field of aviation, and more specifically to a supersonic aircraft. Background Technology

[0002] In the field of civil aviation, increasing the flight speed of civil aircraft is a significant development direction. Compared with conventional subsonic aircraft, supersonic aircraft have higher flight speeds, which can significantly shorten the overall flight time.

[0003] However, existing supersonic aircraft suffer from significant sonic boom problems. Specifically, existing supersonic aircraft generate powerful shock waves (i.e., sonic booms) during supersonic flight. These strong shock waves cause significant noise pollution to people and equipment on the ground, negatively impacting the environment. Furthermore, due to the severe sonic boom problem, the supersonic flight range of existing supersonic aircraft is strictly limited, sometimes only allowing supersonic flight over the sea. As a result, this greatly restricts the route network and coverage of major markets that supersonic aircraft can operate. In short, the sonic boom problem has become a fundamental technical obstacle to commercial supersonic flight over land.

[0004] Furthermore, existing supersonic aircraft are economically unviable. Specifically, the inherent high drag of supersonic aircraft leads to extremely high fuel consumption rates, resulting in significantly higher operating costs compared to subsonic aircraft. Consequently, these high costs result in expensive ticket prices, limiting the potential passenger base and market acceptance, and hindering commercial sustainability.

[0005] Furthermore, existing supersonic aircraft suffer from inadequate takeoff and landing performance. Specifically, the aerodynamic configurations of existing supersonic aircraft are designed to optimize supersonic cruise efficiency, typically resulting in poor lift characteristics at low speeds (especially during takeoff and landing). Consequently, compared to subsonic aircraft, they require longer runways, higher approach speeds, and greater engine thrust. This further increases operational complexity and introduces significant takeoff and landing noise issues.

[0006] Due to the aforementioned problems, supersonic aircraft cannot be commercially operated on a large scale.

[0007] Currently, multiple institutions are conducting research on a new generation of supersonic passenger aircraft. Their main characteristics are low sonic boom pressure distribution during cruise, most of them have slender fuselages that conform to the area law and have large cross-sectional area variations, use large swept wings or delta wings, most of them have conventional or three-surface layouts, use wing-mounted or tail-mounted engines, and have various tail wing forms and installation positions.

[0008] However, the fuselages of these supersonic aircraft are still similar to those of traditional passenger planes, consisting of a nose, a straight section, and a tail. They are not optimized to reduce sonic booms. The shock waves generated by the wings cannot be offset by the expansion waves of the fuselage, and the expansion waves generated by the rapid contraction and lifting of the tail are also difficult to control, ultimately resulting in excessive sonic booms on the ground. Summary of the Invention

[0009] The technical problem that the invention aims to solve

[0010] This application was developed to solve the most critical sonic boom problem among the aforementioned technical issues, and its purpose is to provide a supersonic aircraft capable of achieving a lower sonic boom.

[0011] Technical solutions adopted to solve technical problems

[0012] The first technical solution of this application provides a supersonic aircraft, which includes a fuselage, canards, wings, a tail, and engine nacelles, characterized in that...

[0013] The fuselage is formed in a curved shape such that the curvature line of the fuselage passes through the chord line from top to bottom at least once in the lift direction of the supersonic aircraft, within the range between the first endpoint and the midpoint of the fuselage chord line.

[0014] The camber line is the centerline of the upper and lower profiles of the supersonic aircraft in the lift direction when viewed from the extension direction of the wing.

[0015] The chord is a line segment formed by connecting the foremost and last ends of the supersonic aircraft, which are the first endpoints, in the flight direction.

[0016] According to the first technical solution of this application, a supersonic aircraft can achieve a lower sonic boom. Specifically, by making the fuselage curvature line form a concave curve in the front half of the fuselage (the first 50% of the chord line) (i.e., crossing the chord line from above downwards), the nose of the fuselage generates multiple levels of weak compression waves instead of a single large shock wave, while the curvature shape in the middle and rear can adjust the phase of the expansion wave. This disperses the concentrated high-pressure pulse into multiple low-pressure pulses, thereby significantly reducing the peak value of the sonic boom.

[0017] Based on the first technical solution described above, in the second technical solution of this application, when the length of the chord is set to L, the curvature line passes through the chord at least once from top to bottom within a range of 0.1L~0.3L from the first endpoint to the midpoint along the length direction of the chord.

[0018] According to the supersonic aircraft described in the second technical solution of the present application, by precisely optimizing the shock wave shape at the head of the fuselage, the efficiency of reducing sonic boom can be further improved. Generally speaking, during supersonic flight, the generation position of the initial shock wave at the head (accounting for 60% - 70% of the total sonic boom energy) is basically within the range of the first 20% - 50% of the chord length. Considering this point, by optimizing the shape of the fuselage in this range, further reduction of sonic boom can be achieved.

[0019] Based on the above first technical solution, in the third technical solution of the present application, when the length of the chord line is set as L, the maximum distance between the part of the camber line above the chord line and the chord line is set as C1, and the maximum distance between the part of the camber line below the chord line and the chord line is set as C2, the following relationship is satisfied: 0 < C1 < 0.1L and 0 < C2 < 0.1L.

[0020] According to the supersonic aircraft described in the third technical solution of the present application, while maintaining a significant sonic boom reduction effect, the aerodynamic side effects caused by the bending of the fuselage can be eliminated as much as possible. That is, by limiting the maximum offset of the camber line relative to the chord line, the shock wave distortion / drag surge caused by the bent fuselage can be avoided.

[0021] Based on the above first technical solution, in the fourth technical solution of the present application, the camber line is a piecewise curve or a continuous curve.

[0022] Based on the above fourth technical solution, in the fifth technical solution of the present application, the camber line is one of a spline curve, a Bezier curve, a polynomial curve, and a sine curve.

[0023] According to the supersonic aircraft described in the fourth and fifth technical solutions of the present application, on the premise of meeting low sonic boom, a fuselage design that can be realized in engineering can be provided as much as possible.

[0024] Based on the above first technical solution, in the sixth technical solution of the present application, the engine nacelle is configured to be located between the wing and the tail and partially overlap with the wing when observing the supersonic aircraft from above in the lift direction, and the engine nacelle is also configured to be above the wing in the lift direction.

[0025] Based on the above sixth technical solution, in the seventh technical solution of the present application, when the width of the engine nacelle in the extension direction is set as d and the distance between the leading edge of the engine nacelle and the trailing edge of the wing in the flight direction is set as L0, the following relationship is satisfied: L0 = 0.5d - 2.0d.

[0026] Based on the seventh technical solution mentioned above, in the eighth technical solution of this application, when the distance between the lower surface of the engine nacelle and the upper surface of the wing is set as h when viewed along the extension direction, the following relationship is satisfied: h>0.1d.

[0027] The supersonic aircraft described in the sixth to eighth technical solutions of this application can prevent the shock wave generated at the leading edge of the engine nacelle from propagating downwards.

[0028] Based on the first technical solution described above, in the ninth technical solution of this application, the supersonic aircraft further includes a bulge, which is installed on the lower surface of the fuselage at a position corresponding to the tail fin. When the length of the chord line is set as L, the length of the bulge extending from the tail fin to the engine nacelle along the extension direction is set as L1, and the maximum protrusion height of the bulge from the mounting surface is set as h1, the following relationship is satisfied: 0

[0029] The supersonic aircraft described in the ninth technical solution of this application can eliminate the aerodynamic interference degradation between the tail fin and the engine nacelle, ensuring that the aerodynamic benefits of the low-detonation fuselage design are not offset.

[0030] Based on the ninth technical solution described above, in the tenth technical solution of this application, the supersonic aircraft further includes a ventral fin, which is mounted on a protruding surface of the bulge that protrudes from the mounting surface, and the ventral fin is configured to rotate between a first position and a second position. In the first position, when viewed along the extension direction, the ventral fin protrudes further downward in the lift direction than the protruding surface. In the second position, when viewed along the extension direction, the ventral fin is located on the side of the fuselage.

[0031] The supersonic aircraft described in the tenth technical solution of this application can dynamically adapt to aerodynamic requirements across the entire speed range, eliminating the parasitic sonic boom source of traditional fixed ventral fins during supersonic cruise. Furthermore, it can further reduce sonic boom and improve low-speed directional stability while achieving high-speed drag reduction. Attached Figure Description

[0032] Figure 1 This is a schematic diagram illustrating a supersonic aircraft according to one embodiment of this application.

[0033] Figure 2 This is a side view showing the configuration of the fuselage of a supersonic aircraft according to one embodiment of this application, with other structures omitted.

[0034] Figure 3 This is a top view showing a supersonic aircraft according to one embodiment of this application. ​

[0035] Figure 4 This is a partial side view of a supersonic aircraft according to one embodiment of this application.

[0036] Figure 5 This is another top view showing a supersonic aircraft according to one embodiment of this application.

[0037] Figure 6 This is a front view showing a supersonic aircraft according to one embodiment of this application.

[0038] Figure 7 This is another partial side view showing a supersonic aircraft according to one embodiment of this application.

[0039] Figure 8 This is a pressure distribution diagram showing the area directly beneath the fuselage of a supersonic aircraft according to one embodiment of this application.

[0040] Figure 9 This represents a comparison between the pressure distribution directly beneath the fuselage of a supersonic aircraft according to one embodiment of this application and the pressure distribution directly beneath the fuselage of a conventional supersonic aircraft.

[0041] Symbol Explanation

[0042] S supersonic aircraft

[0043] 1. Fuselage

[0044] 2. Canard

[0045] 3. Wings

[0046] 4. Tail wing

[0047] 5 Engine nacelles

[0048] 6 chords

[0049] 7. Curveline

[0050] 8. Drums

[0051] 9. Pelvic fins

[0052] 10 Pivots Detailed Implementation

[0053] The following is for reference Figures 1 to 7 This application provides a detailed description of a supersonic aircraft according to one embodiment.

[0054] Figure 1 A top view of a supersonic aircraft S according to one embodiment of this application is shown, illustrating the main components of the supersonic aircraft S. For example... Figure 1 As shown, the supersonic aircraft S in this embodiment includes a fuselage 1, canards 2, wings 3, tail 4, and engine nacelles 5.

[0055] The fuselage 1 is an elongated body extending along the flight direction of the supersonic aircraft S, integrating fuel tanks, equipment bays, and passenger cabins. In this application, the fuselage 1 has a curved shape (this will be explained later) and adopts a pointed arched nose cone. Furthermore, the "flight direction" mentioned here refers to the direction of the supersonic aircraft S when it is flying in a straight line at a constant cruise speed at its cruise altitude.

[0056] The canard 2 is symmetrically mounted in the forward region of the fuselage 1. It has aerodynamic functions, used to generate controllable vortices to enhance the lift of the wing 3 (i.e., the main wing). Furthermore, the canard 2 is equipped with, for example, all-moving control surfaces, which provide pitch trim moments to compensate for the forward shift of the aerodynamic focus of the sonic boom fuselage. Figure 5 As shown, the wingspan bc of canard 2 is, for example, 1 / 5 to 1 / 2 of the wingspan b of wing 3.

[0057] The wing 3 is, for example, a large swept wing, with its root transition section located approximately in the middle of the fuselage 1, and has characteristic parameters such as leading and trailing edge sweep angles, aspect ratio, and relative thickness.

[0058] Tail fins 4 are arranged in a pair at the rear of the fuselage 1 in the direction of flight. For example... Figure 6 As shown, when viewed from the nose to the tail along the flight direction of the supersonic aircraft S, the pair of tail fins 4 are arranged in a V-shape, with the tail fins 4 positioned above the engine nacelle 5. The angle θ between the two tail fins 4 can be, for example, 60° to 120°, or arranged in a T-shape.

[0059] The engine nacelle 5 is located between the wing 3 and the tail 4, and can be seen when viewed from the side of the supersonic aircraft S (according to...). Figure 4 As can be seen, the engine nacelle 5 is located above the wing 3. Furthermore, as... Figure 3 As shown, when viewed from above, the engine nacelle 5 partially overlaps with the wing 3.

[0060] Figure 2 A side view of the configuration of the fuselage 1 of a supersonic aircraft S according to an embodiment of this application is shown. For ease of understanding, only the fuselage 1 is shown and other constituent elements are omitted.

[0061] like Figure 2 As shown, the fuselage 1 has a curved shape. Specifically, in the flight direction of the supersonic aircraft S, the fuselage 1 has a convex shape in the front half of its length along the flight direction and a curved shape in the rear half as shown in the figure below. Furthermore, as... Figure 2 As shown, the fuselage 1 is defined by a chord line 6 and a curvature line 7. The chord line 6 is the line segment connecting the foremost point (i.e., the nose) and the rearmost point (i.e., the tail) of the supersonic aircraft in its flight direction. The curvature line 7 is the curve of the supersonic aircraft S. Figure 2 The centerline between the upper and lower contours in the vertical direction (i.e., the lift direction). The "centerline" referred to here means any point on this line (e.g., Figure 2 A line segment is drawn along the lift direction from point P (as shown) and intersects the upper and lower contours of the fuselage 1 (here, small protrusions such as the high-intensity antenna, lights, and fairing) respectively, and the lengths of the resulting line segments satisfy S1=S2. More specifically, the fuselage 1 is formed into a curved shape such that the curvature line 7 of the fuselage 1 passes through the chord 6 from top to bottom at least once within the range between the first endpoint and the midline between the first and second endpoints. Specifically, in this embodiment, the curvature line 7 passes through the chord 6 from top to bottom once within the range of 0.1L~0.3L to 0.5L (i.e., the midpoint) from the first endpoint in the length direction of the chord 6. Here, L represents the length dimension of the chord 6. However, it should be noted that the range of 0.2L~0.5L from the first endpoint is only one example and is not limited to this; it is acceptable as long as it is within the range between the first endpoint and the midpoint. In addition, in Figure 2 The example shown is one in which the curve line 7 crosses the string 6 only once, but it is not limited to this. For example, it is also possible that after crossing the string 6 once, the curve line 7 extends upward and crosses the string 6 again, and then extends downward and crosses the string 6 again. In other words, the curve line 7 can cross the string 6 from top to bottom more than twice.

[0062] In existing supersonic aircraft, the conventional straight cylindrical fuselage generates steep nose compression waves and tail expansion waves during supersonic flight, forming a strong N-shaped pressure wave (the core of the sonic boom), thus producing significant noise. In contrast, the supersonic aircraft according to this embodiment, by designing the fuselage 1 in a curved shape with the curvature line of the fuselage 1 forming a concave (convex) curve in the front half, generates multiple levels of weak compression waves at the nose of the fuselage 1 instead of a single large shock wave. Simultaneously, the curved shape in the middle and rear sections can adjust the phase of the expansion waves. This disperses the concentrated high-pressure pulse into multiple low-pressure pulses, significantly reducing the peak value of the sonic boom. Furthermore, the weak shock wave sequence generated by the fuselage curvature interferes with the shock waves generated by the wings / canards, causing partial phase misalignment between the compression and expansion waves, resulting in a local cancellation effect. This avoids the superposition of shock wave energy, thereby reducing the intensity of the total shock wave energy propagating to the ground. Moreover,

[0063] Furthermore, in this embodiment, optionally, in such a way... Figure 2 When the maximum distance between the portion of the curvature line 7 above the chord line 6 and the chord line 6 is set to C1, and the maximum distance between the portion of the curvature line 7 below the chord line 6 and the chord line 6 is set to C2, the following relationship is satisfied:

[0064] 0 < C1 < 0.1L and 0 < C2 < 0.1.

[0065] By limiting both the maximum distance C1 between the part of the camber line 7 above the chord line 6 and the chord line 6 and the maximum distance C2 between the part of the camber line 7 below the chord line 6 and the chord line 6 within the range of 0 to 0.1L, the negative effects caused by "over-noise reduction" can be avoided. Specifically, if the fuselage 1 is designed to be overly curved, then new shock wave distortions or drag rebounds will occur. When the above maximum distance is greater than 0.1L, the front section of the overly convex fuselage 1 will reflect the expansion wave to form a secondary compression shock wave, which will cause a high-pressure post-pulse in the sonic boom waveform (offsetting the noise reduction benefits in the first half). Therefore, by setting the maximum distances C1 and C2 as above, the monotonicity of the pressure change can be maintained, thus avoiding the negative effects caused by "over-noise reduction".

[0066] In addition, in this embodiment, optionally, the above camber line 7 can be a piecewise curve or a continuous curve. More specifically, the above camber line 7 can be one of a spline curve, a Bezier curve, a polynomial curve, and a sine curve.

[0067] In this embodiment, optionally, as Figure 3 shown, when the width of a single engine nacelle 5 in the extension direction (i.e., spanwise direction) of the wing 3 is set as d and the distance between the leading edge of the engine nacelle 5 and the trailing edge of the wing 3 in the flight direction of the supersonic aircraft S is set as L0, the following relationship is satisfied:

[0068] L0 = 0.5d to 2.0d.

[0069] And, optionally, as Figure 4 shown, when the distance between the lower surface of the engine nacelle 5 and the upper surface of the wing 3 when observed along the extension direction is set as h, the following relationship is satisfied:

[0070] h > 0.1d.

[0071] By arranging the engine nacelle 5 above the wing 3 and making it partially overlap with the wing 3 when observed from above, the前传 shock wave generated by the engine can be blocked by the trailing edge of the wing 3, thereby reducing the ground noise radiation angle. On this basis, by setting the distance between the leading edge of the engine nacelle 5 and the trailing edge of the wing 3 in the flight direction as above, the engine lip shock wave can be embedded in the expansion wave valley of the wing trailing edge, thereby achieving energy neutralization. And by setting the distance h between the lower surface of the engine nacelle 5 and the upper surface of the wing 3 as above, the interaction between the jet flow and the wing surface can be blocked, thereby avoiding the generation of broadband noise.

[0072] In addition, in this embodiment, the supersonic aircraft S may optionally include a bulge 8 and a ventral fin 9.

[0073] like Figure 7 As shown, the bulge 8 is installed on the lower surface of the fuselage 1 at a position corresponding to the tail fin 4. Figure 7 In this design, the length of the bulge 8 extending from the tail 4 to the engine nacelle 5 along the mounting surface of the fuselage 1 when viewed from the side (in other words, along the spanwise direction of the supersonic aircraft S, i.e., the extension direction of the wing 3) is defined as L1, and the maximum protrusion height of the bulge 8 from the aforementioned mounting surface is defined as h1. In this case, the length L of the chord 6, the length L1 of the bulge 8, and the maximum protrusion height h1 of the bulge 8 satisfy the following relationship:

[0074] 0

[0075] Similarly, Figure 7 As shown, the ventral fin 9 is mounted on a protruding surface of the bulge 8 that extends beyond the mounting surface of the fuselage 1. Furthermore, the ventral fin 9 is configured to rotate about a pivot 10 between a first position 9i and a second position 9ii. Specifically, when the supersonic aircraft S is flying at low speed, the ventral fin 9 rotates to the first position 9i. In this first position 9i, when viewed from the side, the ventral fin 9 protrudes further downward than the protruding surface of the bulge 8. This enhances directional stability. On the other hand, when the supersonic aircraft S is cruising at cruising speed, the ventral fin 9 rotates to the second position 9ii. In this second position 9ii, when viewed from the side, the ventral fin 9 is located on the side of the fuselage 1. This reduces drag during supersonic flight.

[0076] Figure 8 This diagram shows the pressure distribution directly beneath the fuselage of the supersonic aircraft S according to this embodiment. Figure 9 The pressure distribution directly beneath the fuselage of the supersonic aircraft S according to this embodiment is shown as a comparison with the pressure distribution directly beneath the fuselage of a conventional supersonic aircraft. Figure 8 and Figure 9 The pressure distribution map is obtained, for example, through computational fluid dynamics simulations or wind tunnel testing. From Figure 9 As can be seen, compared with existing supersonic aircraft, the supersonic aircraft S of this embodiment generates a significantly lower pressure peak during supersonic flight. Therefore, compared with existing supersonic aircraft, the supersonic aircraft S of this embodiment can achieve a reduced sonic boom.

[0077] Furthermore, within the scope of this invention, various embodiments can be freely combined, or appropriately modified or omitted.​

Claims

1. A supersonic aircraft, the supersonic aircraft comprising a fuselage (1), a canard (2), a wing (3), a tail (4) and an engine nacelle (5), characterized in that, the fuselage (1) is formed in a curved shape such that the camber line of the fuselage (1) passes through the chord line from above to below at least once within the range between the first end point and the midpoint of the chord line of the fuselage (1) and in the lift direction of the supersonic aircraft, wherein, the camber line is the midline of the upper contour and the lower contour of the supersonic aircraft in the lift direction when viewed from the extension direction of the wing (3) of the supersonic aircraft, the chord line is a line segment connecting the foremost end as the first end point and the rearmost end of the supersonic aircraft in the flight direction, when the length of the chord line is set as L, the camber line passes through the chord line from above to below at least once within the range from 0.1L to 0.3L from the first end point to the midpoint in the length direction of the chord line.

2. The supersonic aircraft according to claim 1, characterized in that, when the length of the chord line is set as L, the maximum distance between the part of the camber line above the chord line and the chord line is set as C1 and the maximum distance between the part of the camber line below the chord line and the chord line is set as C2, the following relationship is satisfied: 0 < C1 < 0.1L and 0 < C2 < 0.1L.

3. The supersonic aircraft as described in claim 1, characterized in that, The camber line is a piecewise curve or a continuous curve.

4. The supersonic aircraft according to claim 3, characterized in that, the camber line is one of a spline curve, a Bezier curve, a polynomial curve and a sine curve.

5. The supersonic aircraft according to claim 1, characterized in that, the engine nacelle (5) is configured to be located between the wing (3) and the tail (4) and partially overlap with the wing (3) when viewed from above the supersonic aircraft in the lift direction, and the engine nacelle (5) is further configured to be located above the wing (3) in the lift direction.

6. The supersonic aircraft according to claim 5, characterized in that, when the width of the engine nacelle (5) in the extension direction is set as d and the distance between the leading edge of the engine nacelle (5) and the trailing edge of the wing (3) in the flight direction is set as L0, the following relationship is satisfied: L0 = 0.5d to 2.0d.

7. The supersonic aircraft according to claim 6, characterized in that, when the distance between the lower surface of the engine nacelle (5) and the upper surface of the wing (3) when viewed in the extension direction is set as h, the following relationship is satisfied: h > 0.1d.

8. The supersonic aircraft according to claim 1, characterized in that, it further comprises a bulge (8), and the bulge (8) is installed at a position on the lower surface of the fuselage (1) corresponding to the tail (4), When the length of the chord line is set as L, the length of the bulge (8) extending from the tail fin (4) to the engine nacelle (5) in the mounting surface of the bulge (8) and the fuselage (1) when observed along the extension direction is set as L1, and the maximum protrusion height of the bulge (8) protruding from the mounting surface is set as h1, the following relationship is satisfied: 0 < h1 < 0.05L1 and 0 < L1 < 0.15L.

9. The supersonic aircraft according to claim 8, wherein: it further includes a ventral fin (9), the ventral fin (9) is mounted on the protruding surface of the bulge (8) protruding from the mounting surface, and the ventral fin (9) is configured to be rotatable between a first position and a second position, at the first position, when observed along the extension direction, the ventral fin (9) protrudes more downward in the lift direction than the protruding surface, at the second position, when observed along the extension direction, the ventral fin (9) is located on the side surface of the fuselage (1).

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