Target projectile and vehicle

Abstract

The application provides a target projectile and an aircraft. The target projectile comprises a head part in a conical shape, a projectile body comprising at least one engine body section, a first engine body section close to the head part being connected to the head part, a tail part arranged at an end of the engine body section away from the head part and comprising at least two tail wings distributed around an axis of the engine body section, and a rudder part arranged at an end of the first engine body section close to the head part and comprising at least two rudder wings distributed around an axis of the first engine body section, wherein in a first plane perpendicular to the axis of the engine body section, the orthographic projection of the at least two rudder wings is staggered with the orthographic projection of the at least two tail wings. The application realizes controllable supersonic flight of the target projectile in a large airspace and a wide speed range, with a flight height of 0.1-20 km and a flight speed of 0.3-2.4 Ma.

Description

Technical Field

[0001] This application relates to the field of aerospace technology, and more specifically, to a target missile and aircraft. Background Technology

[0002] Supersonic targets generally refer to targets with a Mach number greater than 1 that can simulate the cruise ballistic characteristics of a supersonic cruise aircraft.

[0003] Having a supersonic target missile that can realistically simulate the performance of supersonic cruise aircraft and anti-ship aircraft, or even slightly surpass the simulation, is a basic condition for conducting flight test verification and a prerequisite for in-depth design. Therefore, there is an urgent need for a supersonic cruise target missile that can simulate the ballistic characteristics of supersonic cruise aircraft and anti-ship aircraft.

[0004] In terms of characteristics (such as speed characteristics, ballistic characteristics, etc.), the target missile is very close to the simulated supersonic aircraft. Due to its high price, it is currently mainly used for testing and evaluation and a small amount of combat training. Other target missiles are mainly focused on subsonic and transonic speeds and are difficult to simulate supersonic aircraft.

[0005] Therefore, developing a target missile that can simulate the cruise ballistic characteristics of a supersonic cruise vehicle has great market potential. Summary of the Invention

[0006] This application addresses the shortcomings of existing methods by proposing a target missile and aircraft to solve the technical problems of low target missile performance or high cost in related technologies.

[0007] In a first aspect, embodiments of this application provide a target projectile, comprising:

[0008] The head is cone-shaped;

[0009] The missile body includes at least one engine fuselage section, with the first engine fuselage section near the nose connected to the nose.

[0010] The tail section, located at the end of the engine fuselage section away from the nose, includes at least two tail fins distributed around the axis of the engine fuselage section;

[0011] The control section, located at the end of the first engine fuselage section near the nose, includes at least two control wings distributed around the axis of the first engine fuselage section;

[0012] In this configuration, on a first plane perpendicular to the axis of the engine fuselage section, the orthographic projections of at least two rudder wings are staggered from the orthographic projections of at least two tail fins.

[0013] Optionally, the target projectile includes at least one of the following:

[0014] The control unit includes four control wings, which are evenly distributed around the axis of the first engine fuselage section;

[0015] The tail section includes four tail fins, which are evenly distributed around the axis of the engine fuselage section.

[0016] Optionally, the target projectile includes at least one of the following:

[0017] The four rudder wings of the control unit are arranged in a cross shape on the outer periphery of the first engine fuselage section;

[0018] The four tail fins are arranged in an X-shape around the outer periphery of the engine fuselage section;

[0019] The four rudder wings of the control unit and the four tail wings of the tail section are projected onto the first plane and are alternately and evenly distributed around the axis of the engine fuselage section.

[0020] Optionally, the rudder has:

[0021] The first wing root connects to the outer wall of the first engine fuselage section;

[0022] The first leading edge connects to the first wing root and is positioned close to the head;

[0023] The first trailing edge is connected to the first wing root and is positioned away from the head relative to the first leading edge;

[0024] The first wingtip is connected to the first leading edge and the first trailing edge, respectively;

[0025] The chord length of the first wingtip is greater than the chord length of the first wing root.

[0026] Optionally, the target projectile includes at least one of the following:

[0027] Along a direction away from the axis of the first engine fuselage section, the first leading edge is inclined toward the side closer to the head;

[0028] Along a direction away from the axis of the first engine fuselage section, the first trailing edge is inclined toward the side away from the head;

[0029] The first wingtip has a first front end connected to the first leading edge and a first rear end connected to the first trailing edge. The distance between the first front end and the axis of the first engine fuselage section is less than the distance between the first rear end and the axis of the first engine fuselage section. The first wingtip extends from the first front end to the first rear end.

[0030] Optionally, the target projectile includes at least one of the following:

[0031] The sweep angle of the first leading edge is greater than or equal to 5° and less than or equal to 7°;

[0032] The wingspan of the control unit is greater than or equal to 149.45 mm and less than or equal to 159.45 mm;

[0033] The root-to-tip ratio of the rudder is greater than or equal to 0.52 and less than or equal to 0.56.

[0034] Optionally, the missile body includes a second engine fuselage section connected to the end of the first engine fuselage section away from the nose;

[0035] The target missile also includes at least one of the following:

[0036] At least two tail fins on the first tail section of the first engine fuselage and at least two tail fins on the second tail section of the second engine fuselage are arranged in a one-to-one correspondence with each other on the orthographic projection of the second tail section on the first plane.

[0037] The orthographic projection of the second engine fuselage section onto the first plane overlaps the orthographic projection of the first engine fuselage section onto the first plane;

[0038] The missile body also includes an interstage transition section, through which the second engine fuselage section is connected to the first engine fuselage section.

[0039] Optionally, the length of the second wing root of the tail fin of the first tail section is greater than or equal to 600 mm and less than or equal to 610 mm, the sweep angle of the second leading edge of the tail fin is greater than or equal to 76° and less than or equal to 78°, the length of the second wingtip of the tail fin is greater than or equal to 172 mm and less than or equal to 174 mm, the wingspan of the first tail section is greater than or equal to 92 mm and less than or equal to 100 mm, and the aspect ratio of the first tail section is greater than or equal to 0.14 and less than or equal to 0.16.

[0040] The second tail fin has a second wing root length greater than or equal to 600 mm and less than or equal to 610 mm, a second leading edge sweep angle greater than or equal to 76° and less than or equal to 78°, a second wingtip length greater than or equal to 168 mm and less than or equal to 178 mm, a wingspan greater than or equal to 92 mm and less than or equal to 100 mm, and an aspect ratio greater than or equal to 0.14 and less than or equal to 0.16.

[0041] Optionally, the first end of the interstage transition section near the first engine fuselage section has the same radial dimension as the first engine fuselage section;

[0042] The second end of the interstage transition section, near the second engine fuselage section, has the same radial dimension as the second engine fuselage section;

[0043] The outer wall of the interstage transition section extends from the first end of the interstage transition section to the second end;

[0044] The expansion angle between the outer wall of the interstage transition section and the axis of the engine fuselage section is greater than or equal to 2.87° and less than or equal to 4.87°.

[0045] Secondly, embodiments of this application provide an aircraft that uses the target missile as described above as a target.

[0046] The beneficial technical effects of the technical solutions provided in this application include:

[0047] In this embodiment, the first engine fuselage section of the missile body is connected to the head, and the rudder is located at the end of the first engine fuselage section near the head. Each engine fuselage section of the missile body is provided with a tail fin at the end away from the head. At least two rudder wings of the rudder are distributed around the axis of the first engine fuselage section, and at least two tail fins of the tail fin are distributed around the axis of the engine fuselage section. Furthermore, the orthographic projections of the at least two rudder wings of the rudder on the first plane and the orthographic projections of the at least two tail fins of the tail fin on the first plane are staggered. This arrangement can reduce roll reaction.

[0048] The target missile provided in this application embodiment has a rudder unit installed at the end of the first engine fuselage section near the head, and tail fins installed at the ends of each engine fuselage section away from the head. The rudders of the rudder unit and the tail fins of the tail fins are staggered. The rudder unit's swinging motion completes the three-channel (pitch, yaw, and roll) attitude control of the missile's flight, making the flight attitude controllable throughout the entire flight. Furthermore, due to the shape design of the rudder unit and tail fins, the engine thrust can overcome the flight drag at altitudes from 0.1 to 20 kilometers. Due to the reasonable shape design and low drag, it can achieve flight in a wide airspace range from 0.1 to 20 kilometers. At the same time, through the shape design of the rudders and tail fins, the target missile can meet the speed range of Mach number from 0.3 to 2.4.

[0049] Therefore, the target missile provided in this application embodiment can achieve fully controllable supersonic flight with a flight altitude of 0.1 to 20 kilometers and a flight speed of 0.3 to 2.4 Ma in a wide airspace and a wide speed range.

[0050] Additional aspects and advantages of this application will be set forth in part in the description which follows, and will become apparent from the description or may be learned by practice of this application. Attached Figure Description

[0051] The above and / or additional aspects and advantages of this application will become apparent and readily understood from the following description of the embodiments taken in conjunction with the accompanying drawings, wherein:

[0052] Figure 1 This is a top view of a target projectile provided in an embodiment of this application;

[0053] Figure 2 This is a side view of a target projectile provided in an embodiment of this application;

[0054] Figure 3 This is a schematic diagram of the front view structure of a target projectile provided in an embodiment of this application;

[0055] Figure 4 A side view of the rudder wings, nose, and first engine fuselage section of a target missile provided for an embodiment of this application;

[0056] Figure 5 A side view of the tail fin and second engine fuselage section of a target missile provided in an embodiment of this application;

[0057] Figure 6 A schematic diagram showing the change in rudder deflection angle required for leveling at different Mach numbers during the first stage of flight (angle of attack of -5° and sideslip angle of 0°) provided in an embodiment of this application during target missile launch.

[0058] Figure 7 A schematic diagram showing the change in rudder deflection angle required for leveling at different Mach numbers during the first stage of a target missile launch process (angle of attack of -5° and sideslip angle of 5°) provided in this application embodiment;

[0059] Figure 8 This is a schematic diagram showing the change curve of the rudder deflection angle required for leveling at different Mach numbers during the second stage of a target missile launch process (angle of attack of -5° and sideslip angle of 0°) provided in an embodiment of this application.

[0060] Figure label:

[0061] 100-Target Projectile;

[0062] 10-Head;

[0063] 20-bullet body;

[0064] 21-First engine fuselage section; 22-Second engine fuselage section; 23-Interstage transition section; 24-Cable cover;

[0065] 30 - Tail section;

[0066] 31-Tail fin;

[0067] 311 - Second wing root; 312 - Second leading edge; 313 - Second trailing edge; 314 - Second wing tip;

[0068] 40 - Rudder;

[0069] 41-Rudder;

[0070] 411 - First wing root; 412 - First leading edge; 413 - First trailing edge; 414 - First wing tip. Detailed Implementation

[0071] The embodiments of this application are described below with reference to the accompanying drawings. It should be understood that the embodiments described below with reference to the accompanying drawings are exemplary descriptions for explaining the technical solutions of the embodiments of this application, and do not constitute a limitation on the technical solutions of the embodiments of this application.

[0072] Those skilled in the art will understand that, unless specifically stated otherwise, the terms "described" and "the" as used herein may also include plural forms. It should be further understood that the term "comprising" as used in the specification of this application means the presence of the stated features, integers, and / or components, but does not exclude other features, information, data, components, and / or combinations thereof supported by the art. The term "and / or" as used herein refers to at least one of the items defined by the term; for example, "A and / or B" can be implemented as "A," or as "B," or as "A and B."

[0073] To make the objectives, technical solutions, and advantages of this application clearer, the embodiments of this application will be described in further detail below with reference to the accompanying drawings.

[0074] The target missile and aircraft provided in this application are intended to solve the above-mentioned technical problems in related technologies.

[0075] The technical solution of this application and how it solves the above-mentioned technical problems are described in detail below with specific embodiments. It should be noted that the following embodiments can be referenced, borrowed, or combined with each other, and the same terms, similar features, and similar implementation steps in different embodiments will not be described again.

[0076] This application provides a target projectile 100, the structural schematic diagram of which is shown below. Figures 1 to 3 As shown, it includes: a head 10, a body 20, a tail section 30, and a rudder section 40. The head 10 is conical. The body 20 includes at least one engine fuselage section, and a first engine fuselage section 21 near the head 10 is connected to the head 10. The tail section 30 is located at the end of the engine fuselage section away from the head 10 and includes at least two tail fins 31 distributed around the axis of the engine fuselage section. The rudder section 40 is located at the end of the first engine fuselage section 21 near the head 10 and includes at least two rudder wings 41 distributed around the axis of the first engine fuselage section 21.

[0077] In this configuration, on a first plane perpendicular to the axis of the engine fuselage section, the orthographic projections of at least two rudder wings 41 are staggered from the orthographic projections of at least two tail wings 31.

[0078] In this embodiment, the first engine fuselage section 21 of the missile body 20 is connected to the head 10. The rudder section 40 is located at the end of the first engine fuselage section 21 near the head 10. Each engine fuselage section of the missile body 20 is provided with a tail fin section 30 at the end away from the head 10. At least two rudder wings 41 of the rudder section 40 are distributed around the axis of the first engine fuselage section 21, and at least two tail fins 31 of the tail fin section 30 are distributed around the axis of the engine fuselage section. Furthermore, the orthographic projections of the at least two rudder wings 41 of the rudder section 40 on the first plane are staggered from the orthographic projections of the at least two tail fins 31 of the tail fin section 30 on the first plane. This arrangement can reduce roll reaction.

[0079] The target missile 100 provided in this application embodiment has a rudder 40 installed at the end of the first engine fuselage section 21 of the missile body 20 near the head 10, and a tail fin 30 installed at the end of each engine fuselage section of the missile body 20 away from the head 10. The rudder wings 41 of the rudder 40 and the tail fin 31 of the tail fin 30 are staggered. The three-channel (including pitch, yaw, and roll) attitude control of the missile body 20 is completed by the swing of the rudder 40, so that the flight attitude is controllable throughout the flight. Furthermore, due to the shape design of the rudder 40 and the tail fin 30, the engine thrust can overcome the flight drag at an altitude of 0.1 to 20 kilometers (km). Due to the reasonable shape design and low drag, it can achieve flight in a wide airspace range of 0.1 to 20 km. At the same time, through the shape design of the rudder wings 41 and the tail fin 31, the target missile can meet the speed range of Mach number from 0.3 to 2.4.

[0080] Therefore, the target missile 100 provided in this application embodiment can achieve fully controllable supersonic flight with a flight altitude of 0.1 to 20 km and a flight speed of 0.3 to 2.4 Mach.

[0081] Optionally, such as Figure 1 and Figure 2 As shown in the embodiment of this application, the head 10 includes a head cover. The head cover has a blunt (i.e., the end of the head cover has a small blunt tip, such as a rounded corner) double conical curve. The head cover has a length-to-slenderness ratio of 5±0.1. The head cover length-to-slenderness ratio and the head shape can effectively reduce the drag of the target projectile during flight.

[0082] Optionally, such as Figure 1 and Figure 2 As shown in the embodiment of this application, the head 10 is equipped with a very small lob ball, and the head cover of the payload compartment can use a small-volume double cone to reduce drag.

[0083] Optionally, such as Figures 1 to 3As shown in the embodiment of this application, the rudder section 40 includes four rudder wings 41, which are evenly distributed around the axis of the first engine fuselage section 21. This allows for a uniform load distribution on the rudder section 40, which is beneficial for the smooth operation of the rudder section 40 and improves the stability of the flight attitude control of the target missile 100.

[0084] Optionally, such as Figure 3 As shown in the embodiment of this application, the four rudder wings 41 of the rudder section 40 are arranged in a cross shape on the outer periphery of the first engine fuselage section 21.

[0085] Optionally, such as Figure 3 As shown in the embodiment of this application, of the four control wings 41, one control wing 41 is located at the upper center of the first engine fuselage section 21, denoted as the first control wing (the control deflection angle of the first control wing is denoted by δ1); one control wing 41 is located at the right center of the first engine fuselage section 21, denoted as the second control wing (the control deflection angle of the second control wing is denoted by δ2); one control wing 41 is located at the lower center of the first engine fuselage section 21, denoted as the third control wing (the control deflection angle of the third control wing is denoted by δ3); and one control wing 41 is located at the left center of the first engine fuselage section 21, denoted as the fourth control wing (the control deflection angle of the fourth control wing is denoted by δ4). The four control wings 41 are arranged sequentially in a clockwise direction.

[0086] Optionally, such as Figures 1 to 3 As shown in the embodiment of this application, the tail section 30 includes four tail fins 31, which are evenly distributed around the axis of the engine fuselage section. This allows for a uniform load distribution on the tail section 30, which is beneficial for the stable operation of the tail section 30 and improves the stability of the flight attitude control of the target missile 100.

[0087] Optionally, such as Figure 3 As shown in the embodiment of this application, the four tail wings 31 of the tail wing portion 30 are arranged in an X shape on the outer periphery of the engine fuselage section.

[0088] Optionally, such as Figure 3 As shown in the embodiment of this application, of the four tail fins 31, one tail fin 31 is located at a position 45° to the right of the engine fuselage section, denoted as the first tail fin; one tail fin 31 is located at a position 45° to the right of the engine fuselage section, denoted as the second tail fin; one tail fin 31 is located at a position 45° to the left of the engine fuselage section, denoted as the third tail fin; and one tail fin 31 is located at a position 45° to the left of the engine fuselage section, denoted as the fourth tail fin. The four tail fins 31 are arranged sequentially in a clockwise direction.

[0089] Optionally, such as Figure 3As shown in the embodiment of this application, the orthographic projections of the four rudder wings 41 of the rudder section 40 and the four tail wings 31 of the tail section 30 on the first plane are alternately and evenly distributed around the axis of the engine fuselage section.

[0090] The target missile 100 provided in this application adopts a nose "+" shaped rudder 40 combined with at least one "×" shaped tail fin 30 to achieve three-channel attitude control of at least one stage of engine (each stage of engine corresponds to one engine fuselage section) by a set of rudders 40.

[0091] Optionally, such as Figure 4 As shown in the embodiment of this application, the rudder 41 has a first wing root 411, a first leading edge 412, a first trailing edge 413 and a first wingtip 414.

[0092] The first wing root 411 is connected to the outer wall of the first engine fuselage section 21; the first leading edge 412 is connected to the first wing root 411 and is located close to the nose 10; the first trailing edge 413 is connected to the first wing root 411 and is located away from the nose 10 relative to the first leading edge 412; the first wingtip 414 is connected to the first leading edge 412 and the first trailing edge 413 respectively. The chord length of the first wingtip 414 is greater than the chord length of the first wing root 411.

[0093] Optionally, such as Figure 4 As shown in the embodiment of this application, the first leading edge 412 is inclined toward the side closer to the head 10 along the direction away from the axis of the first engine body section 21.

[0094] Optionally, such as Figure 4 As shown in the embodiment of this application, the first trailing edge 413 is inclined to the side away from the head 10 along the direction away from the axis of the first engine body section 21.

[0095] Optionally, such as Figure 4 As shown in the embodiment of this application, the first wingtip 414 has a first front end connected to the first leading edge 412 and a first rear end connected to the first trailing edge 413. The distance between the first front end and the axis of the first engine fuselage section 21 is less than the distance between the first rear end and the axis of the first engine fuselage section 21. The first wingtip 414 extends from the first front end to the first rear end. Along the direction from the first front end to the first rear end, the outer wall of the first wingtip 414 gradually tilts away from the axis of the first engine fuselage section 21.

[0096] It should be noted that, in the embodiments of this application, the chord length of the first wingtip 414 refers to the length between the first front end and the first rear end of the first wingtip 414, and the chord length of the first wing root 411 refers to the length between the end of the first wing root 411 connected to the first leading edge 412 and the end of the first wing root 411 connected to the first trailing edge 413.

[0097] Optionally, such as Figure 4 As shown in the embodiment of this application, the sweep angle of the first leading edge 412 is greater than or equal to 5° and less than or equal to 7°.

[0098] Optionally, in this embodiment of the application, the wingspan of the rudder 40 is greater than or equal to 149.45 mm and less than or equal to 159.45 mm.

[0099] Optionally, in this embodiment of the application, the root-to-tip ratio of the rudder 41 is greater than or equal to 0.52 and less than or equal to 0.56.

[0100] In this embodiment of the application, by designing the forward sweep angle of the first leading edge 412, the wingspan of the rudder 40, and the root-to-tip ratio of the rudder wing 41, lift and rudder efficiency can be increased, while the servo torque of the rudder 40 can be reduced.

[0101] Specifically, such as Figure 4 As shown in the embodiment of this application, the first leading edge 412 of the rudder 41 has a forward sweep angle of 6°, the wingspan of the rudder 40 is 154.45 mm, and the root-to-tip ratio of the rudder 41 is 0.54 or 0.5405.

[0102] In this embodiment, the chord length of the first wingtip 414 of the rudder 41 is greater than the chord length of the first wing root 411. The rudder 41 adopts a small forward-sweep butterfly wing (or inverted trapezoidal wing) with a forward sweep angle of 6°, which can increase lift and reduce drag to a certain extent. The wingspan is 154.45 mm, and the root-to-tip ratio is 0.54 or 0.5405. The first wingtip 414 is obliquely cut, which can reduce drag and servo control efficiency while ensuring lift and control surface efficiency. The rudder 41 adopts a butterfly wing with a double wedge airfoil, which can provide sufficient lift while facilitating processing and ensuring structural strength. The butterfly rudder shape can reduce the force on the control part 40 during control, thereby reducing the flight drag of the target missile 100 and reducing the range of pressure center variation. The four rudder wings 41 adopt a "+" shaped circumferential layout. The butterfly rudder (i.e., rudder section 40) is located at the front end of the first engine fuselage section 21 (i.e., the end near the head 10). The front rudder 41 and the rear stabilizer wing (i.e., tail wing 31) are staggered to reduce roll reaction.

[0103] Optionally, such as Figure 1 and Figure 2 As shown in the embodiment of this application, the projectile body 20 includes a second engine fuselage section 22 connected to the end of the first engine fuselage section 21 that is away from the head 10.

[0104] In this embodiment, the second engine fuselage section 22 can be a first-stage engine, and the first engine fuselage section 21 can be a second-stage engine.

[0105] Optionally, such as Figures 1 to 3As shown in the embodiment of this application, at least two tail fins 31 of the first tail fin 30 disposed on the first engine fuselage section 21 and at least two tail fins 31 of the second tail fin 30 disposed on the second engine fuselage section 22 are arranged in a one-to-one correspondence with each other on the orthographic projection of the second tail fin 30 on the first plane.

[0106] Specifically, such as Figures 1 to 3 As shown in the embodiment of this application, the rudder section 40 includes four rudder wings 41, which are evenly distributed in a cross shape around the axis of the first engine fuselage section 21, located at the front end of the first engine fuselage section 21 (i.e., the end closer to the head 10), and fixed to the outer periphery of the first engine fuselage section 21. The first tail section 30 includes four tail wings 31, which are evenly distributed in an X shape around the axis of the first engine fuselage section 21, located at the rear end of the first engine fuselage section 21 (i.e., the end farther from the head 10), and fixed to the outer periphery of the first engine fuselage section 21. The second tail section 30 includes four tail wings 31, which are evenly distributed in an X shape around the axis of the second engine fuselage section 22, located at the rear end of the second engine fuselage section 22, and fixed to the outer periphery of the second engine fuselage section 22.

[0107] This application embodiment designs the distribution of the rudder wings 41 of the rudder section 40 and the tail fins 31 of the tail section 30 to obtain a supersonic, fully controllable target missile aerodynamic shape. The target missile 100 with this aerodynamic shape adopts a "+" shaped butterfly rudder at the nose combined with an "×" shaped fixed tail fin for the second-stage engine and an "×" shaped fixed tail fin for the first-stage engine. It can achieve three-channel attitude control for the flight of the two-stage engine with a single butterfly rudder, enabling fully controllable supersonic flight at altitudes from 0.1 to 20 km and speeds from 0.3 to 2.4 Mach.

[0108] Optionally, such as Figure 1 and Figure 2 As shown in the embodiment of this application, the orthographic projection of the second engine fuselage section 22 on the first plane covers the orthographic projection of the first engine fuselage section 21 on the first plane.

[0109] Optionally, such as Figures 1 to 3 As shown in this embodiment, the two-stage engines of the missile body 20 are cylindrical segments with different diameters. The diameter of the second engine fuselage segment 22 corresponding to the first-stage engine is larger than the diameter of the first engine fuselage segment 21 corresponding to the second-stage engine.

[0110] Optionally, such as Figure 5As shown in the embodiment of this application, the tail fin 31 has a second wing root 311, a second leading edge 312, a second trailing edge 313, and a second wingtip 314. The second wing root 311 is connected to the outer wall of the engine fuselage section. The second leading edge 312 is connected to the second wing root 311 and is inclined away from the head 10. The second trailing edge 313 is connected to the second wing root 311. The second wingtip 314 has a second front end and a second rear end, the second front end being connected to the second leading edge 312 and the second rear end being connected to the second trailing edge 313.

[0111] Optionally, such as Figure 1 , Figure 2 and Figure 5 As shown in the embodiment of this application, the length of the second wing root 311 of the tail 31 of the first tail wing portion 30 (i.e., the secondary fixed tail wing) is greater than or equal to 600 mm and less than or equal to 610 mm, the sweep angle of the second leading edge 312 of the tail wing 31 is greater than or equal to 76° and less than or equal to 78°, the length of the second wingtip 314 of the tail wing 31 is greater than or equal to 172 mm and less than or equal to 174 mm, the wingspan of the first tail wing portion 30 is greater than or equal to 92 mm and less than or equal to 100 mm, and the aspect ratio of the first tail wing portion 30 is greater than or equal to 0.14 and less than or equal to 0.16.

[0112] In this embodiment, the above-described configuration provides sufficient lift, reduces drag, and facilitates trim by moving the pressure core aft. A low aspect ratio reduces shape drag and wave drag, while also minimizing rudder reaction, thus ensuring trim control.

[0113] Specifically, such as Figure 1 , Figure 2 and Figure 5 As shown in the embodiment of this application, the second leading edge 312 of the tail fin 31 of the first tail section 30 has a sweep angle of 77° or 77.5°, a wingspan of 96 mm, a second wing root 311 length of 605 mm, and a wingtip (i.e., second wingtip 314) length of 173 mm. The tail fin 31 has a double wedge airfoil with an aspect ratio of 0.15 or 0.1587, which can reduce drag while shifting the pressure center backward. The four tail fins 31 are arranged in an "X" shape at the tail end of the second-stage engine (i.e., the end of the first engine fuselage section 21 away from the nose 10).

[0114] Optionally, such as Figure 1 , Figure 2 and Figure 5As shown in the embodiment of this application, the length of the second wing root 311 of the tail 31 of the second tail wing 30 (i.e., the first-stage fixed tail wing) is greater than or equal to 600 mm and less than or equal to 610 mm, the sweep angle of the second leading edge 312 of the tail wing 31 is greater than or equal to 76° and less than or equal to 78°, the length of the second wingtip 314 of the tail wing 31 is greater than or equal to 168 mm and less than or equal to 178 mm, the wingspan of the second tail wing 30 is greater than or equal to 92 mm and less than or equal to 100 mm, and the aspect ratio of the second tail wing 30 is greater than or equal to 0.14 and less than or equal to 0.16.

[0115] In this embodiment, the above-described configuration provides sufficient lift, reduces drag, and facilitates trim by moving the pressure core aft. A low aspect ratio reduces shape drag and wave drag, while also minimizing rudder reaction, thus ensuring trim control.

[0116] Specifically, such as Figure 1 , Figure 2 and Figure 5 As shown in this embodiment, the second leading edge 312 of the tail fin 31 of the second tail section 30 has a sweep angle of 77° or 77.5°. The sweep is mainly to move the pressure center backward and reduce drag. The wingspan is 96 mm, the length of the second wing root 311 is 605 mm, and the length of the wingtip (i.e., the second wingtip 314) is 173 mm. The tail fin 31 has a double wedge airfoil with an aspect ratio of 0.15 or 0.1587. The four tail fins 31 are arranged in an "X" shape at the tail end of the first-stage engine (i.e., the end of the second engine fuselage section 22 away from the nose 10).

[0117] Optionally, such as Figures 1 to 3 As shown in the embodiment of this application, the missile body 20 further includes an interstage transition section 23, and the second engine fuselage section 22 is connected to the first engine fuselage section 21 through the interstage transition section 23.

[0118] Optionally, such as Figure 1 and Figure 2 As shown in the embodiment of this application, the first end face of the interstage transition section 23 near the first engine fuselage section 21 has the same radial dimension as the first engine fuselage section 21; the second end face of the interstage transition section 23 near the second engine fuselage section 22 has the same radial dimension as the second engine fuselage section 22; the outer wall of the interstage transition section 23 extends from the first end to the second end. The expansion angle between the outer wall of the interstage transition section 23 and the axis of the engine fuselage section is greater than or equal to 2.87° and less than or equal to 4.87°. This arrangement can reduce the wave drag caused by the angle of the interstage transition section 23.

[0119] Specifically, such as Figure 1 and Figure 2As shown in the embodiment of this application, the interstage transition section 23 connects the first-stage engine (i.e., the second engine fuselage section 22) and the second-stage engine (i.e., the first engine fuselage section 21) in series. The interstage transition section 23 adopts a frustum-shaped structure with an expansion angle (also known as a half-cone angle) of 3.87°, which can reduce the wave drag caused by the angle of the interstage transition section 23.

[0120] Optionally, such as Figure 1 and Figure 2 As shown in the embodiment of this application, the first engine block section 21 is cylindrical with a diameter of 215±2mm. The second engine block section 22 is cylindrical with a diameter of 250±2mm. An interstage transition section 23 connects the first engine block section 21 and the second engine block section 22. The end of the interstage transition section 23 connected to the first engine block section 21 has the same diameter as the first engine block section 21, which is 215±2mm. The end of the interstage transition section 23 connected to the second engine block section 22 has the same diameter as the second engine block section 22, which is 250±2mm. The length of the interstage transition section 23 along the axial direction of the engine block section is 260±2mm. The interstage transition section 23 retains a certain length to ensure safe and reliable separation while reducing resistance. Along the first end of the interstage transition section 23 (the end connected to the first engine fuselage section 21), pointing towards the second end of the interstage transition section 23 (the end connected to the second engine fuselage section 22), the diameter of the interstage transition section 23 gradually increases, forming an expanding shape (similar to a horn).

[0121] Optionally, such as Figures 1 to 3 As shown in the embodiment of this application, cable covers 24 for cable cables to pass through are symmetrically distributed on both sides of the first engine body section 21. The cable covers 24 cover the cable cables and protect them.

[0122] Optionally, such as Figure 2 As shown in this embodiment, the head end of the cable cover 24 is designed with a streamlined shape similar to that of a high-speed train to reduce profile drag and wave drag. The cable cover 24 is welded to the first engine fuselage section 21. The cable cover 24 is 1860mm long, 50mm wide, and 15mm thick.

[0123] Optionally, in the embodiments of this application, Figure 6 and Figure 7 A schematic diagram of the curve of the rudder deflection angle δ as a function of Mach number Ma during the flight of the first stage (i.e., the second engine fuselage section 22) of each target missile 100 during the launch process; Figure 8This is a schematic diagram illustrating the variation of the control deflection angle δ with Mach number Ma during the flight of the second stage (i.e., the first engine fuselage section 21) of target missile 100. In this diagram, δ1 represents the control deflection angle of the first control fin, δ2 represents the control deflection angle of the second control fin, δ3 represents the control deflection angle of the third control fin, and δ4 represents the control deflection angle of the fourth control fin. The four control fins 41 are viewed axially from the warhead towards the tail of the missile body 20, with clockwise rotation being positive.

[0124] Specifically, in the embodiments of this application, Figure 6 This shows the rudder deflection at various Mach numbers during the first stage flight of the target missile 100 during launch, with an incoming flow angle of attack α of -5° and an incoming flow sideslip angle β of 0°. Figure 6 It can be determined that during the launch of the target missile 100, the maximum rudder deflection angle required for the three-channel combined leveling of the first-stage engine at various Mach numbers is the rudder deflection angle required for the combined control of pitch, yaw, and roll rudders.

[0125] Figure 7 This shows the rudder deflection at various Mach numbers during the first stage flight of the target missile 100, when the incoming flow angle of attack α is -5° and the incoming flow sideslip angle β is 5°. Figure 7 It can be determined that during the launch of the target missile 100, the maximum rudder deflection angle required for the three-channel combined leveling of the first-stage engine at various Mach numbers is the rudder deflection angle required for the combined control of pitch, yaw, and roll rudders.

[0126] Figure 8 This shows the rudder deflection at various Mach numbers during the second stage of flight of the target missile 100 during launch, when the incoming flow angle of attack α is -5° and the incoming flow sideslip angle β is 0°. Figure 8 It can be determined that during the launch of the target missile 100, the maximum rudder deflection angle required for the three-channel combined leveling of the second-stage engine at various Mach numbers is the rudder deflection angle required for the combined control of pitch, yaw, and roll rudders.

[0127] Optionally, such as Figures 6 to 8 As shown in the embodiments of this application, the flight of the target missile 100 can be balanced under various Mach number conditions.

[0128] according to Figure 6 and Figure 7 It can be seen that the target missile's maximum trim angle at level 100 (corresponding to the second engine fuselage section 22) is -12° during flight, and the maximum rudder deflection angle during trim is -12° (rudder deflection provides a positive pitching moment to the missile body, and a negative deflection angle; the sign of the deflection angle indicates the direction, indicating which direction the rudder is pointing). Based on... Figure 8It can be seen that the maximum trim angle of the target missile 100 second stage (corresponding to the first engine fuselage section 21) during flight is 9.8°. The maximum controllable deflection angle during trim is less than 20° throughout the entire flight. The flight trajectory is within the trimmable angle and Ma range. All data can encompass the flight trajectory, achieving fully controllable flight.

[0129] The target missile 100 provided in this application embodiment achieves drag reduction through the forward-swept leading edge of the butterfly rudder (i.e., rudder section 40), the oblique wingtips, and the swept-back leading edge of the first and second stage stabilizing rudders (i.e., the second tail section 30 and the first tail section 30). Lift is increased by increasing the chord length of the first and second stage stabilizing fins (i.e., tail fins 31). Due to the rational design, drag is low, allowing the engine thrust of the target missile 100 to overcome drag at altitudes from 0.1 to 20 km, enabling flight over a wide airspace range. Simultaneously, the coordinated design of the butterfly rudder and the dimensions of the first and second stage stabilizing fins allows the target missile 100 to cruise within a wide speed range from Mach 0.3 to Mach 2.4, achieving wide-speed-range flight. The target missile 100 uses the oscillation of its front butterfly rudder to control its flight attitude. By oscillating at a certain angle, the target missile 100 can maintain a fixed attitude under certain incoming angle of attack and sideslip angle, achieving fully controllable flight. This enables the target missile 100 to achieve fully controllable flight over a wide airspace and speed range.

[0130] In this embodiment, the butterfly control surface exhibits high control efficiency, low required hinge torque, and fast response characteristics, meeting the needs of new weapon evaluation and combat training. Simultaneously, a single butterfly control system completes two-stage flight three-channel attitude control (achieving a "one-to-two" effect, with the butterfly control system controlling the target missile's first and second stage engines). Compared to transmission technology, this reduces the need for a servo control system, energy equipment, a first-stage trailing edge semi-moving rudder system, and first-stage through-cabin cables and cable covers, thereby reducing the target missile cost by over 100,000 yuan.

[0131] The target projectile provided in this application embodiment can be applied to the field of aerodynamic shape design technology for targets, and further, it can be applied to the aerodynamic shape design of supersonic targets.

[0132] The target missile 100 provided in this application embodiment can be used as a target for simulating various aircraft characteristics (such as speed characteristics, altitude characteristics, maneuverability, etc.).

[0133] Based on the same inventive concept, this application provides an aircraft that uses the target missile 100 as described above as a target.

[0134] It should be noted that since the aircraft provided in this application embodiment uses the target missile 100 provided in this application embodiment as the target, the aircraft provided in this application embodiment also has the above-mentioned beneficial effects of the target missile 100 provided in this application embodiment, which will not be repeated here.

[0135] Optionally, in the embodiments of this application, the aircraft can be a weapon system, including but not limited to anti-ship weapons (e.g., anti-ship missiles), air defense weapons (e.g., air defense missiles, anti-aircraft guns, etc.), aviation weapons (e.g., air-to-air missiles, etc.), and missile weapons (e.g., hypersonic missiles, cruise missiles, etc.).

[0136] By applying the embodiments of this application, at least the following beneficial effects can be achieved:

[0137] The target missile provided in this application achieves drag reduction through a design that incorporates a forward-swept leading edge of the butterfly rudder (i.e., control fin), sloping wingtips, and a swept-back leading edge of the first and second stage stabilizing fins (i.e., the second and first tail fins). Lift is increased by enhancing the chord length of the first and second stage stabilizing fins (i.e., tail fins). Due to the rational design, drag is low, allowing the missile's engine thrust to overcome drag at altitudes from 0.1 to 20 km, enabling flight over a wide airspace. Simultaneously, the coordinated design of the butterfly rudder and the first and second stage stabilizing fins allows for cruise flight over a wide speed range from Mach 0.3 to Mach 2.4. The missile uses the swinging of the butterfly rudder to control its flight attitude; swinging at a certain angle allows the missile to maintain a fixed attitude at specific angles of attack and sideslip angles, achieving fully controllable flight. This enables the missile to achieve fully controllable flight over a wide airspace and speed range.

[0138] In the description of this application, the terms "center," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate directions or positional relationships based on the exemplary directions or positional relationships shown in the accompanying drawings. They are used to facilitate the description or simplification of the embodiments of this application and are not intended to indicate or imply that the device or component 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 this application.

[0139] The terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this application, unless otherwise stated, "a plurality of" means two or more.

[0140] In the description of this application, it should be noted that, unless otherwise expressly specified and limited, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal communication between two components. Those skilled in the art can understand the specific meaning of the above terms in this application based on the specific circumstances.

[0141] In the description of this specification, specific features, structures, materials, or characteristics may be combined in any suitable manner in one or more embodiments or examples.

[0142] The above description is only a partial implementation of this application. It should be noted that for those skilled in the art, other similar implementation methods based on the technical concept of this application, without departing from the technical concept of this application, also fall within the protection scope of the embodiments of this application.

Claims

1. A target projectile, characterized in that, include: The head is cone-shaped; The projectile body includes at least one engine fuselage section, with a first engine fuselage section near the head connected to the head; the projectile body also includes a second engine fuselage section connected to the end of the first engine fuselage section away from the head. The tail section, located at the end of the engine fuselage section away from the head, includes at least two tail fins distributed around the axis of the engine fuselage section; The rudder section, located at one end of the first engine fuselage section near the nose, includes at least two rudder wings distributed around the axis of the first engine fuselage section; Wherein, on a first plane perpendicular to the axis of the engine fuselage section, the orthographic projections of the at least two rudder wings are staggered from the orthographic projections of the at least two tail wings; At least two tail fins disposed on the first tail section of the first engine fuselage and at least two tail fins disposed on the second tail section of the second engine fuselage are arranged in a one-to-one correspondence with each other on the orthographic projection of the second tail section on the first plane. The rudder has: The first wing root is connected to the outer wall of the first engine fuselage section; The first leading edge is connected to the first wing root and is positioned close to the head; The first trailing edge is connected to the first wing root and is disposed away from the head relative to the first leading edge; The first wingtip is connected to the first leading edge and the first trailing edge, respectively; The chord length of the first wingtip is greater than the chord length of the first wing root; The first wingtip has a first front end connected to the first leading edge and a first rear end connected to the first trailing edge. Along the direction from the first front end to the first rear end, the outer wall of the first wingtip gradually tilts away from the axis of the first engine fuselage section.

2. The target projectile according to claim 1, characterized in that, Includes at least one of the following: The rudder section includes four rudder wings, which are evenly distributed around the axis of the first engine fuselage section. The tail section includes four tail fins, which are evenly distributed around the axis of the engine fuselage section.

3. The target projectile according to claim 2, characterized in that, Includes at least one of the following: The four rudder wings of the rudder are arranged in a cross shape on the outer periphery of the first engine fuselage section; The four tail fins of the tail section are arranged in an X-shape on the outer periphery of the engine fuselage section; The orthographic projections of the four rudder wings of the rudder section and the four tail wings of the tail section onto the first plane are alternately and evenly distributed around the axis of the engine fuselage section.

4. The target projectile according to claim 1, characterized in that, Includes at least one of the following: Along a direction away from the axis of the first engine fuselage section, the first leading edge is inclined toward the side closer to the head; Along a direction away from the axis of the first engine fuselage section, the first trailing edge is inclined toward a side away from the head; The distance between the first front end and the axis of the first engine fuselage section is less than the distance between the first rear end and the axis of the first engine fuselage section, and the first wingtip extends from the first front end to the first rear end.

5. The target projectile according to claim 1, characterized in that, Includes at least one of the following: The sweep angle of the first leading edge is greater than or equal to 5° and less than or equal to 7°; The wingspan of the rudder is greater than or equal to 149.45 mm and less than or equal to 159.45 mm; The root-to-tip ratio of the rudder is greater than or equal to 0.52 and less than or equal to 0.

56.

6. The target projectile according to any one of claims 1 to 3, characterized in that, The target project also includes at least one of the following: The orthographic projection of the second engine fuselage section onto the first plane covers the orthographic projection of the first engine fuselage section onto the first plane; The missile body also includes an interstage transition section, and the second engine fuselage section is connected to the first engine fuselage section through the interstage transition section.

7. The target projectile according to claim 6, characterized in that, The length of the second wing root of the first tail fin is greater than or equal to 600 mm and less than or equal to 610 mm, the sweep angle of the second leading edge of the tail fin is greater than or equal to 76° and less than or equal to 78°, the length of the second wingtip of the tail fin is greater than or equal to 172 mm and less than or equal to 174 mm, the wingspan of the first tail fin is greater than or equal to 92 mm and less than or equal to 100 mm, and the aspect ratio of the first tail fin is greater than or equal to 0.14 and less than or equal to 0.

16. The second tail fin has a second wing root length greater than or equal to 600 mm and less than or equal to 610 mm, a second leading edge sweep angle greater than or equal to 76° and less than or equal to 78°, a second wingtip length greater than or equal to 168 mm and less than or equal to 178 mm, a wingspan greater than or equal to 92 mm and less than or equal to 100 mm, and an aspect ratio greater than or equal to 0.14 and less than or equal to 0.

16.

8. The target projectile according to claim 6, characterized in that, The first end of the interstage transition section near the first engine fuselage section has the same radial dimension as the first engine fuselage section; The second end of the interstage transition section near the second engine fuselage section has the same radial dimension as the second engine fuselage section; The outer wall of the interstage transition section extends from the first end to the second end of the interstage transition section; The expansion angle between the outer wall of the interstage transition section and the axis of the engine fuselage section is greater than or equal to 2.87° and less than or equal to 4.87°.

9. An aircraft, characterized in that, The target is a target projectile as described in any one of claims 1 to 8.

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