A supersonic civil aircraft layout

Abstract

The application belongs to the field of aviation technology, and particularly relates to a supersonic civil aircraft layout. The supersonic civil aircraft layout of the application is characterized in that an engine nacelle is installed above the rear fuselage between a wing and a horizontal tail, an air inlet is shielded by a wing trailing edge below, an air outlet is shielded by a horizontal tail leading edge below, shock waves generated by the engine are prevented from spreading to the ground by the shielding effect of the wing and the horizontal tail, and the ground sound blast loudness of the aircraft is effectively reduced; the engine and the pylon can be deflected by a small angle around a rotating shaft in the pylon to adjust the installation angle of the engine and improve the propulsion efficiency of the engine under non-design point working conditions; the wing tip of the horizontal tail can be folded upward to provide additional vertical tail capacity when the aircraft cruises at supersonic speed, compensate for the loss of vertical tail efficiency when the aircraft cruises at supersonic speed, and prevent the spread of expansion waves generated by the engine tail nozzle and the horizontal tail to the ground; the wing tip is lowered when the aircraft takes off and lands at low speed to ensure a large horizontal tail area to meet the trim requirements of the take-off and landing configuration.

Description

Technical Field

[0001] The present application belongs to the field of aviation technology, and in particular relates to a supersonic civil aircraft layout. Background Art

[0002] In the design process of supersonic civil aircraft layout, it is necessary to focus on the sonic boom suppression and the balance between engine propulsion efficiency in the full speed range, as well as the reduction in vertical tail efficiency during supersonic cruise.

[0003] When a conventional supersonic civil aircraft is cruising at supersonic speed, the shock waves generated by the engine's intake and exhaust systems merge with the shock waves generated by other parts of the aircraft. When transmitted to the ground, they greatly increase the loudness of the ground sonic boom, bringing significant difficulties to the airworthiness certification of such aircraft.

[0004] In current supersonic civil aircraft research being conducted by countries and related research institutions around the world, engines and their intake and exhaust systems are generally designed for maximum flight Mach numbers. Operating under conditions deviating from this design point can lead to rapid performance degradation, resulting in a series of adverse consequences, including reduced engine propulsion efficiency, increased fuel consumption, and increased external resistance. To avoid the impact of sonic booms on the daily lives of people on the ground when flying over densely populated urban areas, supersonic civil aircraft typically employ subsonic cruise. At this point, the engine deviates from its design point, reducing propulsion efficiency.

[0005] During supersonic cruise, the tail capacity of the vertical tail of a supersonic civil aircraft will drop significantly, and the heading stability will also drop.

[0006] Therefore, it is desired to have a technical solution to overcome or at least alleviate at least one of the above-mentioned deficiencies of the prior art. Summary of the Invention

[0007] The purpose of this application is to provide a supersonic civil aircraft layout to solve at least one problem existing in the prior art.

[0008] The technical solution of this application is:

[0009] A supersonic civil aircraft layout, comprising:

[0010] body;

[0011] a left wing located on the left side of the center rear of the fuselage;

[0012] a right wing located on the right side of the center rear of the fuselage;

[0013] a left trailing edge flap located inboard of the left wing trailing edge;

[0014] a right trailing edge flap located inboard of the right wing trailing edge;

[0015] a left trailing edge aileron, located outside the trailing edge of the left wing;

[0016] a right trailing edge aileron, located outboard of the trailing edge of the right wing;

[0017] a left engine nacelle, mounted on the fuselage behind the left wing via a left engine nacelle pylon;

[0018] a right engine nacelle, mounted on the fuselage behind the right wing via a right engine nacelle pylon;

[0019] a tail-mounted left engine, mounted in the left engine nacelle;

[0020] a tail-mounted right engine, mounted in the right engine nacelle;

[0021] A fully movable horizontal stabilizer, comprising a left horizontal stabilizer and a right horizontal stabilizer, wherein the left horizontal stabilizer is located on the left side of the tail of the fuselage, and the right horizontal stabilizer is located on the right side of the tail of the fuselage;

[0022] The vertical tail is located on the upper side of the tail of the fuselage, and a rudder is installed on the trailing edge of the vertical tail.

[0023] In at least one embodiment of the present application,

[0024] The connection between the front end of the left wing and the fuselage is located at 1 / 2 the length of the fuselage;

[0025] The front end of the right wing is connected to the fuselage at a position 1 / 2 the length of the fuselage.

[0026] In at least one embodiment of the present application,

[0027] The left wing has a wing root leading edge sweep angle of 75° and a wing tip leading edge sweep angle of 60°;

[0028] The right wing has a wing root leading edge sweep angle of 75° and a wing tip leading edge sweep angle of 60°.

[0029] In at least one embodiment of the present application,

[0030] The curvature of the wing root leading edge and the wing tip leading edge of the left wing are smoothly transitioned;

[0031] The curvature of the wing root leading edge and the wing tip leading edge of the right wing are smoothly transitioned.

[0032] In at least one embodiment of the present application,

[0033] The left engine nacelle pylon is connected to the fuselage via a left engine nacelle pylon rotating shaft;

[0034] The right pylon of the engine nacelle is connected to the fuselage via a right pylon rotating shaft of the engine nacelle.

[0035] In at least one embodiment of the present application,

[0036] The left nacelle pylon rotation axis is located at 3 / 5 of the chord length at the root and 1 / 3 of the chord length at the tip of the left nacelle pylon, and the left nacelle pylon can pitch and deflect around the left nacelle pylon rotation axis;

[0037] The right pylon rotation axis of the engine nacelle is located at 3 / 5 of the chord length at the root and 1 / 3 of the chord length at the tip of the right pylon of the engine nacelle. The right pylon of the engine nacelle can pitch and deflect around the right pylon rotation axis of the engine nacelle.

[0038] In at least one embodiment of the present application,

[0039] The left engine air inlet of the left engine nacelle is located above the trailing edge of the left wing;

[0040] The right engine air intake of the right engine nacelle is located above the right wing trailing edge.

[0041] In at least one embodiment of the present application,

[0042] The left engine exhaust port of the left engine nacelle is located above the leading edge of the left horizontal tail;

[0043] The right engine exhaust port of the right engine nacelle is located above the leading edge of the right horizontal tail.

[0044] In at least one embodiment of the present application,

[0045] The deflection angle range of the left horizontal tail is +20° to -15°;

[0046] The deflection angle range of the right horizontal tail is +20° to -15°.

[0047] In at least one embodiment of the present application,

[0048] A foldable horizontal tail wingtip portion is provided at the wingtip of the left horizontal tail via a left horizontal tail wingtip rotation shaft;

[0049] A foldable horizontal tail wingtip portion is provided at the wingtip of the right horizontal tail via a right horizontal tail wingtip rotating shaft.

[0050] In at least one embodiment of the present application,

[0051] The left horizontal tail wing tip rotation axis is arranged at the wing tip of the left horizontal tail at 85% of the span;

[0052] The right horizontal tail wing tip rotation axis is arranged at a wing tip at 85% of the span of the right horizontal tail.

[0053] In at least one embodiment of the present application, the foldable portion of the horizontal tail wing tip can be folded upward by 90°.

[0054] In at least one embodiment of the present application,

[0055] The left horizontal tail is a truncated delta wing;

[0056] The right horizontal tail is a pointed delta wing.

[0057] In at least one embodiment of the present application,

[0058] The leading edge sweep angle of the left horizontal tail is 60°;

[0059] The leading edge sweep angle of the right horizontal tail is 60°.

[0060] In at least one embodiment of the present application, the trailing edge of the vertical tail is located at a 7 / 8 chord length position of the all-moving horizontal tail.

[0061] In at least one embodiment of the present application, the leading edge sweep angle of the vertical tail is 54°.

[0062] The invention has at least the following beneficial technical effects:

[0063] In the supersonic civil aircraft layout of the present application, the engine nacelle is installed above the rear fuselage between the wing and the horizontal tail, with the lower part of the air intake shielded by the trailing edge of the wing, and the lower part of the exhaust shielded by the leading edge of the horizontal tail. The shielding effect of the wing and the horizontal tail is used to prevent the shock wave generated by the engine from propagating to the ground, effectively reducing the ground sonic boom loudness of this type of aircraft; the engine and the pylon can be deflected in a small angle of pitch around the rotating axis in the pylon, so as to adjust the installation angle of the engine and improve the propulsion efficiency of the engine under non-design point conditions; the horizontal tail wingtip can be folded upward during supersonic cruise to provide additional vertical tail tail capacity, compensate for the loss of vertical tail efficiency during supersonic cruise, and prevent the expansion wave generated by the engine tail nozzle and horizontal tail from propagating to the ground; during low-speed takeoff and landing, the wingtip is lowered to ensure a larger horizontal tail area to meet the takeoff and landing configuration balancing requirements. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] Figure 1 This is a side view of the layout of a supersonic civil aircraft according to one embodiment of the present application;

[0065] Figure 2 This is a top view of the layout of a supersonic civil aircraft according to one embodiment of the present application;

[0066] Figure 3 This is a front view of the layout of a supersonic civil aircraft according to one embodiment of the present application;

[0067] Figure 4 This is an axial view of the supersonic civil aircraft layout according to one embodiment of the present application.

[0068] in:

[0069] 1-fuselage; 2-vertical tail; 3-rudder; 4-right wing; 5-left wing; 6-right engine air intake; 7-left engine air intake; 8-right nacelle pylon shaft; 9-left nacelle pylon shaft; 10-right nacelle pylon; 11-left nacelle pylon; 12-right engine exhaust port; 13-left engine exhaust port; 14-right horizontal tail wingtip shaft; 15-left horizontal tail wingtip shaft; 16-left engine nacelle; 17-right engine nacelle; 18-foldable horizontal tail wingtip portion. DETAILED DESCRIPTION

[0070] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the drawings in the embodiments of this application. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described embodiments are part of the embodiments of this application, not all of the embodiments. The embodiments described below with reference to the drawings are exemplary and are intended to be used to explain this application, and should not be understood as limitations on this application. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The embodiments of this application are described in detail below in conjunction with the drawings.

[0071] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating the orientation or position relationship, are based on the orientation or position relationship shown in the accompanying drawings, and are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation. Therefore, it should not be understood as limiting the scope of protection of this application.

[0072] The following is combined with Figures 1 to 4 This application is described in further detail.

[0073] The present application provides a supersonic civil aircraft layout, including: a fuselage 1, a left wing 5, a right wing 4, a left trailing edge flap, a right trailing edge flap, a left trailing edge aileron, a right trailing edge aileron, a left engine nacelle 16, a right engine nacelle 17, a tail-mounted left engine, a tail-mounted right engine, a fully movable horizontal tail, and a vertical tail 2.

[0074] Specifically, the left wing 5 is located at the left side of the center and rear of the fuselage 1; the right wing 4 is located at the right side of the center and rear of the fuselage 1; the left trailing edge flap is located on the inner side of the trailing edge of the left wing 5; the right trailing edge flap is located on the inner side of the trailing edge of the right wing 4; the left trailing edge aileron is located on the outer side of the trailing edge of the left wing 5; and the right trailing edge aileron is located on the outer side of the trailing edge of the right wing 4.

[0075] In a preferred embodiment of the present application, the front end of the left wing 5 connects to the fuselage 1 at a position 1½ the length of the fuselage; the front end of the right wing 4 connects to the fuselage 1 at a position 1½ the length of the fuselage. The left wing 5 has a root leading edge sweep angle of 75° and a wingtip leading edge sweep angle of 60°; the right wing 4 has a root leading edge sweep angle of 75° and a wingtip leading edge sweep angle of 60°. The left wing 5 has a smooth transition between the root and wingtip leading edges; the right wing 4 has a smooth transition between the root and wingtip leading edges.

[0076] The left engine nacelle 16 is installed on the fuselage 1 behind the left wing 5 through the engine nacelle left pylon 11; the right engine nacelle 17 is installed on the fuselage 1 behind the right wing 4 through the engine nacelle right pylon 10; the tail-mounted left engine is installed in the left engine nacelle 16; and the tail-mounted right engine is installed in the right engine nacelle 17.

[0077] In a preferred embodiment of the present application, the left nacelle pylon 11 is connected to the fuselage 1 via a left nacelle pylon pivot shaft 9; the right nacelle pylon 10 is connected to the fuselage 1 via a right nacelle pylon pivot shaft 8. The left nacelle pylon pivot shaft 9 is located at the base of the left nacelle pylon 11, 3 / 5 of the chord length, and the tip of the left nacelle pylon 11, 1 / 3 of the chord length. The left nacelle pylon 11 is capable of pitching and yawing about the left nacelle pylon pivot shaft 9. The right nacelle pylon pivot shaft 8 is located at the base of the right nacelle pylon 10, 3 / 5 of the chord length, and the tip of the right nacelle pylon 10, 1 / 3 of the chord length. The right nacelle pylon 10 is capable of pitching and yawing about the right nacelle pylon pivot shaft 8. The left engine air intake 7 of the left nacelle 16 is located above the trailing edge of the left wing 5; the right engine air intake 6 of the right nacelle 17 is located above the trailing edge of the right wing 4. The left engine exhaust port 13 of the left engine nacelle 16 is located above the leading edge of the left horizontal tail; the right engine exhaust port 12 of the right engine nacelle 17 is located above the leading edge of the right horizontal tail.

[0078] The all-movable horizontal stabilizer includes a left horizontal stabilizer and a right horizontal stabilizer. The left horizontal stabilizer is located on the left side of the tail of the fuselage 1, and the right horizontal stabilizer is located on the right side of the tail of the fuselage 1; the vertical stabilizer 2 is located on the upper side of the tail of the fuselage 1, and a rudder 3 is installed on the trailing edge of the vertical stabilizer 2.

[0079] In a preferred embodiment of the present application, the deflection angle range of the left horizontal stabilizer is +20° to -15°; the deflection angle range of the right horizontal stabilizer is +20° to -15°. A foldable horizontal stabilizer wingtip portion 18 is provided at the wingtip of the left horizontal stabilizer via a left horizontal stabilizer wingtip rotation axis 15; a foldable horizontal stabilizer wingtip portion 18 is provided at the wingtip of the right horizontal stabilizer via a right horizontal stabilizer wingtip rotation axis 14. The left horizontal stabilizer wingtip rotation axis 15 is provided at the wingtip at 85% of the span of the left horizontal stabilizer; the right horizontal stabilizer wingtip rotation axis 14 is provided at the wingtip at 85% of the span of the right horizontal stabilizer. The foldable horizontal stabilizer wingtip portion 18 can be folded upward by 90°. The left horizontal stabilizer is a truncated delta wing; the right horizontal stabilizer is a truncated delta wing. The leading edge sweep angle of the left horizontal stabilizer is 60°; the leading edge sweep angle of the right horizontal stabilizer is 60°. The trailing edge of the vertical stabilizer 2 is located at the 7 / 8 chord length of the full-moving horizontal stabilizer. The leading edge sweep angle of the vertical tail 2 is 54°.

[0080] Advantageously, the layout of this supersonic civil aircraft, through the arrangement of the engine air inlet and exhaust ports, allows the high-temperature, high-speed airflow of the engine tail jet to be blocked by the upper surface of the all-moving horizontal tail, and the spanwise position of the tail jet is blocked by the upward-folded all-moving horizontal tail wing tip.

[0081] The supersonic civil aircraft layout of this application can solve the following problems existing in the use of supersonic civil aircraft:

[0082] (1) During supersonic cruise, the shock waves / expansion waves generated by the engine intake / exhaust system merge with the shock waves / expansion waves generated by other aircraft components and, when transmitted to the ground, greatly increase the loudness of the ground sonic boom;

[0083] (2) In the cruising state of a supersonic civil aircraft, the engine thrust forms an angle with the incoming airflow due to the aircraft's own angle of attack, the nacelle installation angle, and the downward deflection angle of the engine tail nozzle. That is, the thrust axis has a small positive angle of attack relative to the incoming airflow. This angle results in a loss of thrust in the direction of flight, but an increase in lift in the direction of flight. Therefore, for a fixed lift-to-drag ratio, there should theoretically be an optimal thrust axis deflection.

[0084] The deflection of the thrust axis is directly related to the aircraft's lift-to-drag ratio. Assuming the aircraft's lift-to-drag ratio is K, and the angle between the thrust axis and the direction of flight is 0, the thrust is T, and the lift is KT. When the thrust axis is at an angle α with the direction of flight, the thrust T has an increment in the lift direction, Tsinα. Therefore, the actual lift of the aircraft is reduced to KT - Tsinα. From this, it can be deduced that the required thrust in the direction of flight should be T - Tsinα / K, while the actual thrust is Tcosα. Therefore, the residual thrust is Tcosα - (T - Tsinα / K). A greater residual thrust indicates higher propulsion efficiency, so the problem is transformed into finding the maximum value of the following equation.

[0085] Tcosα-(T-Tsinα / K)

[0086] Assuming T and K to be constants, and making the derivative of this formula with respect to α equal to 0, we can get the value of α when the above formula reaches its maximum value.

[0087] The maximum propulsion efficiency is derived as follows:

[0088] α=arctan(1 / K)

[0089] Supersonic civil aircraft engines and their intake and exhaust systems are generally designed with the supersonic cruise flight Mach number as their design point. During subsonic cruise, the lift-to-drag ratio K SB Deviating from its supersonic cruise lift-to-drag ratio K SP , the engine deviates from its supersonic cruise design point and the propulsion efficiency is reduced.

[0090] (3) During supersonic cruise, the efficiency of the vertical tail of a supersonic civil aircraft will decrease significantly, and the heading stability will also decrease. In order to meet the heading stability requirements during supersonic cruise, the vertical tail tail capacity and vertical tail area are usually designed to be larger, and the structural weight increases accordingly.

[0091] The supersonic civil aircraft layout of the present application provides a low sonic boom supersonic civil aircraft layout. By designing a fully movable horizontal tail with wing tips that can be folded upward by 90 degrees, the horizontal tail wing tips fold upward during supersonic cruise, providing additional vertical tail tail capacity, compensating for the loss of vertical tail efficiency during supersonic cruise, and at the same time preventing the expansion waves generated by the engine tail nozzle and horizontal tail from propagating to the ground, thereby reducing the loudness of the cruise ground sonic boom. By designing an engine pylon that can perform a small angle pitch deflection around its rotation axis, the engine installation angle during subsonic cruise is adjusted to arctan(1 / K SB ), so that its residual thrust is maximized in this state and its cruising efficiency is highest; by designing the air intake to be located in front of the upper part of the wing trailing edge and the exhaust to be located in the engine nacelle above the leading edge of the horizontal tail, the shielding effect of the wing and horizontal tail is used to prevent the shock wave generated by the engine from propagating to the ground, effectively reducing the contribution of the ground sonic boom caused by the engine and its intake and exhaust system.

[0092] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A supersonic civil aircraft layout, characterized in that: include: fuselage (1); A left wing (5) is located on the left side of the rear center of the fuselage (1); a right wing (4), located on the right side of the center rear of the fuselage (1); A left trailing edge flap located inside the trailing edge of the left wing (5); a right trailing edge flap located inboard of the trailing edge of the right wing (4); A left trailing edge aileron, located outside the trailing edge of the left wing (5); a right trailing edge aileron located outside the trailing edge of the right wing (4); A left engine nacelle (16) is mounted on the fuselage (1) behind the left wing (5) via a left engine nacelle hanger (11); A right engine nacelle (17) is mounted on the fuselage (1) behind the right wing (4) via a right engine nacelle hanger (10); A tail-mounted left engine, mounted in the left engine nacelle (16); A tail-mounted right engine, mounted in the right engine nacelle (17); A fully movable horizontal tail, comprising a left horizontal tail and a right horizontal tail, wherein the left horizontal tail is located on the left side of the tail of the fuselage (1), and the right horizontal tail is located on the right side of the tail of the fuselage (1); A foldable wingtip portion (18) is provided at the wingtip of the left horizontal tail via a left horizontal tail wingtip rotation shaft (15); A foldable wingtip portion (18) is provided at the wingtip of the right horizontal tail via a right horizontal tail wingtip rotation shaft (14); A vertical tail (2) is located on the upper side of the tail of the fuselage (1), and a rudder (3) is installed on the trailing edge of the vertical tail (2).

2. The supersonic civil aircraft layout according to claim 1, characterized in that: The front end of the left wing (5) is connected to the fuselage (1) and is located at a position 1 / 2 the length of the fuselage (1); The connection between the front end of the right wing (4) and the fuselage (1) is located at a position 1 / 2 the length of the fuselage (1).

3. The supersonic civil aircraft layout according to claim 2, characterized in that: The left wing (5) has a wing root leading edge sweep angle of 75° and a wing tip leading edge sweep angle of 60°; The right wing (4) has a wing root leading edge sweep angle of 75° and a wing tip leading edge sweep angle of 60°.

4. The supersonic civil aircraft layout according to claim 3, characterized in that: The curvature of the wing root leading edge and the wing tip leading edge of the left wing (5) are smoothly transitioned; The curvature of the wing root leading edge and the wing tip leading edge of the right wing (4) are smoothly transitioned.

5. The supersonic civil aircraft layout according to claim 1, characterized in that: The engine nacelle left pylon (11) is connected to the fuselage (1) via the engine nacelle left pylon rotating shaft (9); The engine nacelle right pylon (10) is connected to the fuselage (1) via the engine nacelle right pylon rotating shaft (8).

6. The supersonic civil aircraft layout according to claim 5, characterized in that: The engine nacelle left pylon rotation axis (9) is located at a position of 3 / 5 of the chord length at the root and 1 / 3 of the chord length at the tip of the engine nacelle left pylon (11), and the engine nacelle left pylon (11) can perform pitch deflection around the engine nacelle left pylon rotation axis (9); The engine nacelle right pylon rotation axis (8) is located at a position of 3 / 5 of the chord length at the root and 1 / 3 of the chord length at the tip of the engine nacelle right pylon (10), and the engine nacelle right pylon (10) can perform pitch deflection around the engine nacelle right pylon rotation axis (8).

7. The supersonic civil aircraft layout according to claim 6, characterized in that: The left engine air inlet (7) of the left engine nacelle (16) is located above the trailing edge of the left wing (5); The right engine air inlet (6) of the right engine nacelle (17) is located above the trailing edge of the right wing (4).

8. The supersonic civil aircraft layout according to claim 7, characterized in that: The left engine exhaust port (13) of the left engine nacelle (16) is located above the leading edge of the left horizontal tail; The right engine exhaust port (12) of the right engine nacelle (17) is located above the leading edge of the right horizontal tail.

9. The supersonic civil aircraft layout according to claim 1, characterized in that: The deflection angle range of the left horizontal tail is +20° to -15°; The deflection angle range of the right horizontal tail is +20° to -15°.

10. The supersonic civil aircraft layout according to claim 9, characterized in that: The left horizontal tail wing tip rotation shaft (15) is arranged at the wing tip of the left horizontal tail at 85% of the span; The right horizontal tail wing tip rotation shaft (14) is arranged at a wing tip at 85% of the span of the right horizontal tail.

11. The supersonic civil aircraft layout according to claim 10, characterized in that: The horizontal tail wing tip foldable portion (18) can be folded upwards by 90 degrees.

12. The supersonic civil aircraft layout according to claim 10, characterized in that: The left horizontal tail is a truncated delta wing; The right horizontal tail is a pointed delta wing.

13. The supersonic civil aircraft layout according to claim 12, characterized in that: The leading edge sweep angle of the left horizontal tail is 60°; The leading edge sweep angle of the right horizontal tail is 60°.

14. The supersonic civil aircraft layout according to claim 13, characterized in that: The trailing edge of the vertical tail (2) is located at a 7 / 8 chord length position of the fully movable horizontal tail.

15. The supersonic civil aircraft layout according to claim 14, characterized in that: The leading edge sweep angle of the vertical tail (2) is 54°.

Images