CN118293753A

CN118293753A - Carrier rocket and attitude control method thereof

Info

Publication number: CN118293753A
Application number: CN202410358929.5A
Authority: CN
Inventors: 梁纪秋; 胡长伟; 杜林霏; 张鹏飞; 彭威; 熊闯; 陈腾芳; 张健鹏; 刘宪闯
Original assignee: General Designing Institute of Hubei Space Technology Academy
Current assignee: General Designing Institute of Hubei Space Technology Academy
Filing date: 2024-03-27
Publication date: 2024-07-05

Abstract

The invention relates to the technical field of aerospace, in particular to a carrier rocket and a gesture control method thereof. The carrier rocket comprises: a fairing, a load, an N-stage engine and an attitude control mechanism. Wherein the fairing comprises a fairing front cover and a fairing rear cover; the load is arranged in the rear cover of the fairing; the 1 st to N-2 nd engines in the N-level engines are arranged in the forward direction and are sequentially connected with the fairing back cover, and the N-1 st engine and the N-th engine are arranged in a flip-chip manner and are arranged in the fairing front cover; the attitude control mechanism is arranged at the circumferential outer edge of the fairing front cover and is used for controlling the forward flight attitude of the whole carrier rocket and the 180-degree reverse direction of the rest part after the fairing rear cover is separated. The scheme can solve the problems of the prior art that the integral rigidity and strength performance of the rocket are reduced and the stable control quality is poor due to the expansion of the fairing envelope and the increase of the rocket length, and has high capacity envelope and high control quality.

Description

Carrier rocket and attitude control method thereof

Technical Field

The invention relates to the technical field of aerospace, in particular to a carrier rocket and a gesture control method thereof.

Background

Along with the expansion of future aerospace activities, the envelope size of the spacecraft has a development trend of large scale under the same mass level.

Therefore, the carrier rocket basic stage length and the loading capacity are properly increased, and the fairing size is enlarged, so that the capacity elasticity of the rocket can be improved, and the rocket can adapt to the launching tasks of more spacecrafts under the same carrying capacity. In view of convenience and suitability for transportation, an increase in arrow length is generally considered.

However, the increase of the length of the whole rocket and the expansion of the size of the fairing can lead to overlarge slenderness ratio of the rocket, reduced overall rigidity and strength performance, and difficult bearing transportation and flight overload; meanwhile, the large-size fairing causes the forward movement of the whole rocket core, the large pneumatic interference and the deterioration of stable control quality, and the overall layout design of the rocket is difficult to achieve both high-capacity envelope and high control quality.

Disclosure of Invention

Aiming at the defects in the prior art, the invention aims to provide a carrier rocket and a gesture control method thereof, which can solve the problems of reduced overall rigidity and strength performance and poor stable control quality of the rocket caused by expanding the envelope of a fairing and increasing the length of the rocket in the prior art.

In order to achieve the above purpose, the invention adopts the following technical scheme:

in one aspect, the present invention provides a launch vehicle comprising:

A fairing comprising a fairing front shroud and a fairing rear shroud;

A load disposed within the cowling aft cowl;

The N-stage engine, wherein the 1 st to N-2 nd stage engines are arranged in the forward direction and are sequentially connected with the fairing back cover, and the N-1 st stage engine and the N th stage engine are arranged in a flip-chip manner and are arranged in the fairing front cover;

the attitude control mechanism is arranged at the circumferential outer edge of the fairing front cover and is used for controlling the forward flight attitude of the whole carrier rocket and the 180-degree reverse direction of the rest part of the fairing rear cover after the fairing rear cover is separated.

In some alternatives, the attitude control mechanism includes a first attitude control assembly and a second attitude control assembly disposed at intervals along the fairing axial direction, the first attitude control assembly being located at an end remote from the fairing aft cowl, wherein the first attitude control assembly includes: four first engine groups of group that are located on the first circular cross-section of radome fairing, four groups first engine groups are the cross setting, and all are located radome fairing outward flange department, second gesture control assembly includes: the four groups of second engine units are arranged on the second circular section of the fairing, the four groups of second engine units are arranged in an X shape and are all positioned at the outer edge of the fairing, the first circular section and the second circular section are perpendicular to the axial direction of the fairing, the coordinate systems with the same orientation are established on the first circular section and the second circular section, the normal line of the nozzle opening of the first engine unit is in the same direction as the axis of the coordinate system, the four groups of second engine units are respectively positioned in four quadrants of the coordinate system, the two pairs of second engine units are symmetrical relative to the axis of the coordinate system, and the normal line of the nozzle opening is in an angle with the axis of the coordinate system.

In some alternatives, the fairing front shroud includes first and second diverging cones of different taper spliced in sequence, the second diverging cone being adjacent to a side of the fairing rear shroud.

In some alternatives, the outer profile curve of the fairing satisfies von karman curve equation.

In some alternative solutions, the front cowling and the rear cowling are connected through a transition cabin, the nozzle of the nth stage engine is located in the second expansion cone, the nozzle of the nth stage engine is located in the first expansion cone, the nth stage engine and the load are connected with the transition cabin, and the nth stage engine is connected with the transition cabin through a connection structure.

In some alternative schemes, the fairing back cover comprises a cylindrical section and an inverted cone section which are sequentially connected, the cylindrical section is connected with the transition connecting section, and an included angle between the inverted cone section and the axis direction of the cylindrical section is 19-21 degrees.

On the other hand, the invention also provides a carrier rocket attitude control method for controlling the flying attitude of any carrier rocket, which comprises the following steps:

The 1 st to N-2 nd engines are utilized to push the carrier rocket to fly to a set height or to be separated after flying for a set time, and the cowling back cover is separated;

controlling the rest part to reverse 180 degrees after the rear cover of the fairing is separated by using a gesture control mechanism;

and the front cowling of the fairing is thrown off, and the N-1 stage engine and the N stage engine are utilized to continue to advance the load.

In some alternative schemes, when the flying attitude of the whole carrier rocket is controlled, the pitching maneuver or the yawing maneuver is performed through the first attitude control component, the second attitude control component is used as a backup, and when the rolling maneuver is performed, the second attitude control component is used for controlling.

In some alternatives, the control of the attitude of the flight segment prior to rudder detachment is performed:

The control moment in the pitching direction is as follows:

Wherein l ₁ is the axial distance from the thrust line of the first engine set to the center of mass of the arrow body, F ₁ is the thrust of the first engine set, l ₂ is the axial distance from the thrust line of the second engine set to the center of mass of the arrow body, F ₂ is the thrust of the second engine set, Pitch control moment provided for the rudder; Is rudder deflection angle of The pneumatic pitching moment coefficient of the rocket is q is dynamic pressure, S is pneumatic reference area, L is pneumatic reference length,Is rudder deflection angle ofThe aerodynamic pitching moment coefficients of the rocket are u1 to u8, which are respectively on-off control instructions of the T1 to T8 engine units, wherein 1 is on or 0,1 is off, and 0 is on;

the control moment of the yaw channel is as follows:

Wherein M _ψ0＝qSLC_n(δ_ψ0), a yaw control moment provided for the rudder; c _n(δ_ψ0) is rudder deflection angle The pneumatic yaw moment coefficient of the rocket;

the control moment of the rolling channel is as follows:

Wherein M _γ0＝qSLC_l(δ_γ0) provides a roll control moment for the air rudder, and h ₁ is the distance from the mounting surface of the first engine unit (41) to the longitudinal axis of the rocket body; c _l(δ_γ0) is a rocket aerodynamic roll moment coefficient when the rudder deflection angle is delta _γ0;

When the attitude control of the flight section after the rudder is separated is performed:

The control moment in the pitching direction is as follows:

the control moment of the yaw channel is as follows:

the control moment of the rolling channel is as follows:

In some alternatives, the controlling, by the attitude control mechanism, 180 degrees reversal of the remaining portion of the cowling after the cowling is detached includes: and controlling 180-degree reversal of the rest part of the fairing after the fairing is separated by using four first engine groups of the first attitude control assembly, and performing attitude control and backup by using four second engine groups of the second attitude control assembly.

Compared with the prior art, the invention has the advantages that: and the N-1 level engine and the N level engine are reversely arranged in the fairing front cover, after the 1 st to N-2 level engines and the fairing rear cover are separated, the rest part of the fairing rear cover is reversed by 180 degrees after the fairing rear cover is separated by utilizing the gesture control mechanism at the circumferential outer edge of the fairing front cover, so that the N-1 level engine and the N level engine continue to push loads to advance. The space utilization rate of the head part of the fairing is greatly improved, the length of a full arrow is shortened, the relatively reasonable slenderness ratio is obtained, and the negative quality of a cabin structure is reduced; and the carrying coefficient is improved, and the unit effective load transmitting cost is effectively reduced.

Drawings

In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for the description of the embodiments will be briefly described below, and it is apparent that the drawings in the following description are only some embodiments of the present application, and other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.

FIG. 1 is a schematic view of a carrier rocket according to an embodiment of the present invention;

FIG. 2 is a schematic view of section A-A of FIG. 1 in accordance with an embodiment of the present invention;

FIG. 3 is a schematic view of section B-B of FIG. 1 in accordance with an embodiment of the invention;

FIG. 4 is a schematic view of a cowling back cover according to an embodiment of the present invention;

FIG. 5 is a schematic view of a rear cowling of a cowling according to an embodiment of the present invention;

FIG. 6 is a schematic view of a front cover of a fairing removed in accordance with an embodiment of the invention.

In the figure: 1. a fairing; 11. a fairing front shroud; 111. a first expansion cone; 112. a second expansion cone; 12. a cowling back cover; 121. a cylindrical section; 122. a back taper section; 2. load; 31. an N-1 st stage engine; 32. an nth stage engine; 33. a rocket base stage; 4. a posture control mechanism; 41. a first engine block; 42. a second engine block; 5. a transition cabin section; 6. an air rudder; 7. and a connection structure.

Detailed Description

For the purpose of making the objects, technical solutions and advantages of the embodiments of the present application more apparent, the technical solutions of the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present application, and it is apparent that the described embodiments are some embodiments of the present application, but not all embodiments of the present application. All other embodiments, which can be made by those skilled in the art based on the embodiments of the application without making any inventive effort, are intended to be within the scope of the application.

Embodiments of the present invention are described in further detail below with reference to the accompanying drawings.

As shown in fig. 1, the present invention provides a launch vehicle, comprising: a fairing 1, a load 2, an N-stage engine and an attitude control mechanism.

Wherein the fairing 1 comprises a fairing front shroud 11 and a fairing rear shroud 12; the load 2 is arranged in the fairing back cover 12; the 1 st to N-2 nd engines of the N-stage engines are arranged in the forward direction and are sequentially connected with the fairing back cover 12, and the N-1 st engine 31 and the N-th engine 32 are arranged in a flip-chip manner and are arranged in the fairing front cover 11; the attitude control mechanism is arranged at the circumferential outer edge of the fairing front cover 11 and is used for controlling the forward flight attitude of the whole carrier rocket and the 180-degree reverse direction of the rest part after the fairing rear cover 12 is separated.

When the carrier rocket is launched, the carrier rocket is pushed to fly to a set height or separated after flying for a set time by sequentially utilizing the 1 st to N-2 nd engines, and the fairing back cover 12 is separated; the rest 180 degrees of the rear cover 12 of the fairing are reversed after being separated by the gesture control mechanism; the cowl front 11 is thrown off, and the load 2 is continuously propelled forward by the N-1 stage engine 31 and the N stage engine 32. In the scheme, the N-1 stage engine 31 and the N stage engine 32 are inversely installed in the fairing front cover 11, after the 1 st to N-2 stage engines and the fairing rear cover 12 are separated, the rest part of the fairing rear cover 12 separated by the fairing rear cover 12 is 180 degrees reversed by utilizing the attitude control mechanism at the circumferential outer edge of the fairing front cover 11, so that the N-1 stage engine 31 and the N stage engine 32 continue to push the load 2 to advance. The space utilization rate of the head part of the fairing is greatly improved, the length of a full arrow is shortened, the relatively reasonable slenderness ratio is obtained, and the negative quality of a cabin structure is reduced; and the carrying coefficient is improved, and the unit effective load transmitting cost is effectively reduced.

Referring to fig. 2 and 3, in some alternative embodiments, the attitude control mechanism 4 includes a first attitude control assembly and a second attitude control assembly disposed at intervals along the axial direction of the fairing 1, the first attitude control assembly being located at an end remote from the fairing aft cowl 12, wherein the first attitude control assembly includes: four first engine groups 41 on the first circular cross section of the fairing 1, the four first engine groups 41 are arranged in a cross shape and are all located at the outer edge of the fairing 1, and the second attitude control assembly comprises: the four groups of second engine units 42 are arranged on the second circular section of the fairing 1, the four groups of second engine units 42 are arranged in an X shape and are all arranged at the outer edge of the fairing 1, the first circular section and the second circular section are perpendicular to the axial direction of the fairing 1, a coordinate system with the same orientation is established on the first circular section and the second circular section, the normal line of the nozzle opening of the first engine unit 41 is in the same direction as the coordinate system line, the four groups of second engine units 42 are respectively arranged in four quadrants of the coordinate system, and the normal line of the nozzle opening is in an angle of 45 degrees with the axis of the coordinate system.

In this embodiment, the first attitude control components and the second attitude control components are respectively disposed on two circular sections of the fairing 1 that are disposed at intervals, four first engine units 41 of the first attitude control components are disposed at intervals of 90 degrees in the circumferential direction of the first circular section, and the nozzle opening normal line of the first engine units 41, that is, the nozzle opening direction is in the same direction as the axis of the coordinate system and faces away from the circle point of the coordinate system, such as T1, T2, T3 and T4 in fig. 2, are used to control the deflection in four directions; as shown in fig. 3, four second engine blocks 42 of the second attitude control assembly are disposed in an X-shape on the first circular cross section, and are respectively located in four quadrants of the coordinate system, that is, each second engine block 42 is disposed between two adjacent first engine blocks 41. The nozzle opening normals of the second engine block 42, that is, the nozzle opening directions of the second engine block 42 are at an angle of 45 degrees to a coordinate system axis, the four sets of the second engine blocks 42 are symmetrical with respect to the coordinate system axis, and the nozzle opening directions are directed away from the coordinate system dots, such as T5, T6, T7 and T8 in fig. 3.

The first attitude control assembly is positioned at one end far away from the fairing back cover 12, namely a first circular section is close to the arrow tip and far away from the mass center of the whole rocket, the diameter of a second circular section is larger than that of the first circular section, and the first engine units 41 are arranged on the first circular section, so that the four groups of first engine units 41 can control the pitching and deflecting actions of the whole rocket more easily due to larger deflecting moment applied to the whole rocket; because the diameter of the second circular section is larger than that of the first circular section, four groups of second engine units 42 are arranged on the second circular section, so that the axial distance between the second engine units 42 and the whole rocket is larger, the rolling moment is easier to be applied to the rocket, and the rolling gesture of the whole rocket is convenient to control; in addition, the nozzle opening direction of the second engine unit 42 forms an angle of 45 degrees with an axis of a coordinate system, and can be used as a backup of the first engine unit 41 in a pairwise matching manner, and when the first engine unit 41 fails, the second engine unit 42 can also control pitch and yaw of the rocket in a pairwise matching manner.

In this example, the 1 st to N-2 nd engines are collectively referred to as rocket base stage 33, and 4 air rudders 6 are mounted at the tail section of rocket base stage 33.

When the flight direction and the attitude of the carrier rocket are controlled by the first attitude control assembly and the second attitude control assembly:

In the pitch passage, when the positive pitch control is performed: then T3 works and T6, T7 back-ups; negative pitch control is performed: t1 works, and T5 and T8 are backed up; the rudder 6 is optionally engaged.

In the yaw path: when the positive yaw control is performed: t2 works, and T5 and T6 are backed up; when negative yaw control is performed: t4 works, and T7 and T8 are backed up; the rudder 6 is optionally engaged.

In the roll channel: when forward rolling control is performed: t6 and T8 work; when negative roll control is performed: t5 and T7 work; the rudder 6 is optionally engaged.

In this example, the first engine unit 41 and the second engine unit 42 are both composed of 1 large-thrust attitude control engine and 1 small-thrust attitude control engine, and the on-off of the different thrust engines is controlled according to the real-time interference moment in the flying process, so that the control accuracy degree can be improved while the stable control capability is ensured, and the consumption of the propellant can be reduced.

In some alternative embodiments, the fairing front shroud 11 includes a first flared cone 111 and a second flared cone 112 that sequentially blend in different tapers, the second flared cone 112 being adjacent to one side of the fairing rear shroud 12.

In this embodiment, the N-1 stage engine 31 and the N-1 stage engine 32 are disposed in the storage spaces of the first expansion cone 111 and the second expansion cone 112, the engine nozzle of the N-1 stage engine 31 is accommodated in the first expansion cone 11 and is adapted to the storage space inside the first expansion cone 11, and the volume of the N-1 stage engine 32 is smaller, and by accommodating the engine nozzle of the N-1 stage engine 31 in the first expansion cone 11, the storage space utilization of the entire fairing front cover 11 can be improved.

In some alternative embodiments, the outer contour curve of the fairing 1 satisfies the von karman curve equation.

The von karman curve equation is as follows:

θ＝cos^-1(1-2x/L)

Wherein θ is an intermediate variable, L is a von karman curve length, that is, a length of the front hood 11 of the fairing, and x and y are an abscissa and an ordinate of the von karman curve in a rectangular coordinate system, respectively; r is the maximum radius of the cowl front 11. The two-section fold line formed by three points on the von Karman curve is used as a two-cone-section bus of the front cover of the fairing, as shown in the following figure 1, wherein the length ratio L ₁/L₂ of the two cones is 1.1-1.2.

In some alternative embodiments, as shown in fig. 4, the fairing front hood 11 and the fairing rear hood 12 are connected by the transition piece 5, the nozzle of the nth stage engine 32 is positioned in the second expansion cone 112, the nozzle of the nth-1 stage engine 31 is positioned in the first expansion cone 111, the nth stage engine 32 and the load 2 are connected to the transition piece 5, and the nth-1 stage engine 31 is connected to the transition piece 5 by the connection structure 7.

In this embodiment, the N-1 stage engine 31 is connected to the transition piece 5 by the connection structure 7 and accommodates the receiving space within the first expansion cone 111 and the second expansion cone 112. The load 2 is connected with the transition cabin section 5 through the rocket adapter and is accommodated in the accommodating space of the second expansion cone 112. The connection structure 7 in this example is a rod system structure.

In some alternative embodiments, the fairing section 12 includes a cylindrical section 121 and an inverted cone section 122 connected in sequence, where the cylindrical section 121 is connected to the transition section 321, and the inverted cone section 122 has an angle of 19 ° to 21 ° with respect to the axial direction of the cylindrical section 121.

In this embodiment, the included angle between the inverted cone section 122 and the axial direction of the cylindrical section 121 is 19-21 degrees, so that the aerodynamic drag and the overall slenderness ratio of the whole rocket can be optimized, and the cylindrical section 121 and the inverted cone section 122 with the contracted shape are integrally formed.

On the other hand, the invention provides a carrier rocket attitude control method for controlling the flying attitude of any carrier rocket, which comprises the following steps:

referring to fig. 4, S1: the engines of stages 1 to N-2 are used for pushing the carrier rocket to fly to a set height or to fly for a set time to be separated, and the fairing back cover 12 is separated.

In this example, the thrust of the first engine group 41 defining the mounting surface where the first circular section is located is F1, and the axial distances from the thrust line to the center of mass of the arrow body are l1 respectively; the thrust of a single engine on a mounting surface where the second circular section is positioned on the fairing front cover 1 is defined as F2, the distance from the mounting surface to the mass center of the arrow body is l2, and the distance from the mounting surface to the longitudinal axis of the arrow body is h1. The attitude control engine layout is shown in fig. 1.

In this example, the launch vehicle includes a 4-stage engine.

The flight section attitude control before the rudder 6 is disengaged includes: the combined control method of the first attitude control component, the second attitude control component and the tail section air rudder 6 is adopted, wherein the first attitude control component, the second attitude control component and the tail section air rudder 6 are respectively arranged on two circular sections of the front cowling 11 at intervals. And determining whether the first attitude control component and the second attitude control component are started or not and whether the air rudder 6 participates in control or not according to the attitude control requirements in the flight process. A typical joint control scheme is as follows: in the initial take-off section, as the control moment provided by the air rudder is too small, the pitching and yawing control is carried out by a first gesture control component in the gesture control mechanism of the arrow head; and a second gesture control component in the arrow head gesture control mechanism performs rolling control and forms a pitching yaw control backup. After t1, the roll control shift rudder is carried out, the second engine unit 42 of the second attitude control assembly stops working, and the first attitude control assembly in the arrow head attitude control mechanism continuously controls pitching yaw; after time t2, the first motor unit 41 of the first attitude control assembly is completely stopped, and the air rudder controls three channels of pitching yaw roll. U1 to u8 are defined as on-off control instructions of the T1 to T8 engine units respectively, wherein 1 is on or 0,1 is off, and 0 is on.

In some alternative embodiments, the control of the attitude of the entire launch vehicle is performed by the first attitude control assembly during pitch maneuvers or yaw maneuvers, and the second attitude control assembly is used as a backup during roll maneuvers.

Specifically, in the pitch passage, when the positive pitch control is performed: then T3 works and T6, T7 back-ups; negative pitch control is performed: t1 works, and T5 and T8 are backed up; the rudder 6 is optionally engaged.

The control moment in the pitching direction is as follows:

Wherein, The pitch control moment provided for the rudder, q is dynamic pressure, S is pneumatic reference area, L is pneumatic reference length,Is rudder deflection angle ofRocket aerodynamic pitching moment coefficient.

The control moment of the yaw channel is as follows:

Wherein M _ψ0＝qSLC_n(δ_ψ0), a yaw control moment provided for the rudder;

c _n(δ_ψ0) is rudder deflection angle The aerodynamic yaw moment coefficient of the rocket.

The control moment of the rolling channel is as follows:

Wherein, M _γ0＝qSLC_l(δ_γ0) provides a rolling control moment for the air rudder, and h ₁ is the distance from the mounting surface of the first engine unit 41 to the longitudinal axis of the rocket body;

C _l(δ_γ0) is the rocket aerodynamic roll moment coefficient at rudder deflection angle delta _γ0.

And (3) attitude control of the flight section after the rudder 6 is separated: the combined control method of the first attitude control assembly and the second attitude control assembly is adopted, wherein the first attitude control assembly and the second attitude control assembly are respectively arranged on two circular sections of the front fairing 11 at intervals.

In the pitch passage, when the positive pitch control is performed: then T3 works and T6, T7 back-ups; negative pitch control is performed: t1 works, and T5 and T8 are backed up.

In the yaw path: when the positive yaw control is performed: t2 works, and T5 and T6 are backed up; when negative yaw control is performed: t4 works, and T7 and T8 are backed up.

In the roll channel: when forward rolling control is performed: t6 and T8 work; when negative roll control is performed: t5 and T7 work.

The control moment in the pitching direction is as follows:

the control moment of the yaw channel is as follows:

the control moment of the rolling channel is as follows:

Referring to fig. 5, S2: the remaining 180 degrees of the fairing after the aft cowl 12 is disengaged are reversed by an attitude control mechanism.

In this example, after the 1 st to N-2 nd stage engines and the cowl top cover 12 are separated, the remaining portion is reversed 180 degrees after the cowl top cover 12 is separated by using the attitude control mechanism at the circumferential outer edge of the cowl top cover 11. Specifically, four first engine blocks 41 of the first attitude control assembly are used to control 180-degree reversal of the remaining portion of the cowling back 12 after it is disengaged, and attitude control and backup are performed by four second engine blocks 42 of the second attitude control assembly.

Referring to fig. 6, S3: the cowl front 11 is thrown off, and the load 2 is continuously propelled forward by the N-1 stage engine 31 and the N stage engine 32.

In this example, after 180 degrees reversing the remainder of the rocket base stage 33 and fairing 12 after being thrown off, the fairing front shroud 11 is thrown off and the load 2 is propelled forward to a predetermined altitude or orbit by the N-1 stage engine 31 and the N-th stage engine 32.

In summary, the present embodiment inverts the N-1 st stage engine 31 and the N-th stage engine 32 in the cowl front cover 11, and after the 1 st to N-2 nd stage engines and the cowl rear cover 12 are separated, the remaining portions 180 degrees reverse after the cowl rear cover 12 is separated by the posture control mechanism at the outer circumferential edge of the cowl front cover 11, so that the N-1 st stage engine 31 and the N-th stage engine 32 continue to propel the load 2. The space utilization rate of the head part of the fairing is greatly improved, the length of a full arrow is shortened, the relatively reasonable slenderness ratio is obtained, and the negative quality of a cabin structure is reduced; and the carrying coefficient is improved, and the unit effective load transmitting cost is effectively reduced. The front cover of the fairing adopts a biconical aerodynamic shape similar to a von Karman curve, and the structure can reduce aerodynamic resistance and production process difficulty while ensuring volume ratio.

In the description of the present application, it should be noted that the azimuth or positional relationship indicated by the terms "upper", "lower", etc. are based on the azimuth or positional relationship shown in the drawings, and are merely for convenience of describing the present application and simplifying the description, and are not indicative or implying that the apparatus or element in question must have a specific azimuth, be constructed and operated in a specific azimuth, and thus should not be construed as limiting the present application. Unless specifically stated or limited otherwise, the terms "mounted," "connected," and "coupled" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present application can be understood by those of ordinary skill in the art according to the specific circumstances.

It should be noted that in the present application, relational terms such as "first" and "second" and the like are used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Moreover, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising one … …" does not exclude the presence of other like elements in a process, method, article, or apparatus that comprises the element.

The foregoing is only a specific embodiment of the application to enable those skilled in the art to understand or practice the application. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the application. Thus, the present application is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A launch vehicle, comprising:

a fairing (1) comprising a fairing front shroud (11) and a fairing rear shroud (12);

A load (2) provided in the cowling back cover (12);

The N-stage engine, wherein the 1 st to N-2 nd stage engines are arranged in the forward direction and are sequentially connected with the fairing back cover (12), and the N-1 st stage engine (31) and the N-th stage engine (32) are arranged in a flip-chip manner and are arranged in the fairing front cover (11);

the attitude control mechanism (4) is arranged at the circumferential outer edge of the fairing front cover (11) and is used for controlling the forward flight attitude of the whole carrier rocket and the 180-degree reverse direction of the rest part after the fairing rear cover (12) is separated.

2. A launch vehicle according to claim 1, wherein the attitude control mechanism (4) comprises a first attitude control assembly and a second attitude control assembly arranged at intervals along the axial direction of the fairing (1), the first attitude control assembly being located at an end remote from the fairing aft cowl (12), wherein the first attitude control assembly comprises: four groups of first engine groups (41) on the first circular section of the fairing (1), wherein the four groups of first engine groups (41) are arranged in a cross shape and are all positioned at the outer edge of the fairing (1), and the second attitude control assembly comprises: the four groups of second engine units (42) are arranged on the second circular section of the fairing (1), the four groups of second engine units (42) are arranged in an X shape and are all arranged at the outer edge of the fairing (1), the first circular section and the second circular section are perpendicular to the axial direction of the fairing (1), coordinate systems with the same orientation are established on the first circular section and the second circular section, the normal line of a nozzle opening of the first engine unit (41) is in the same direction as the axis of the coordinate system, the four groups of second engine units (42) are respectively arranged in four quadrants of the coordinate system, the two pairs of second engine units are symmetrical relative to the axis of the coordinate system, and the normal line of the nozzle opening is in an angle of 45 degrees with the axis of the coordinate system.

3. A launch vehicle according to claim 1, wherein the fairing front shroud (11) comprises a first expansion cone (111) and a second expansion cone (112) of different taper, spliced in sequence, the second expansion cone (112) being adjacent to one side of the fairing back shroud (12).

4. A launch vehicle according to claim 3, characterized in that the outer profile curve of the fairing (1) satisfies the von karman curve equation.

5. A launch vehicle according to claim 3, wherein the fairing front cowling (11) and the fairing rear cowling (12) are connected by a transition piece (5), the nozzle of the nth stage engine (32) is located in the second expansion cone (112), the nozzle of the nth-1 stage engine (31) is located in the first expansion cone (111), the nth stage engine (32) and the load (2) are connected with the transition piece (5), and the nth-1 stage engine (31) is connected with the transition piece (5) by a connection structure (7).

6. A launch vehicle according to claim 5, wherein the fairing back shroud (12) comprises a cylindrical section (121) and an inverted cone section (122) which are connected in sequence, the cylindrical section (121) is connected with the transition connecting section (321), and an included angle between the inverted cone section (122) and the axis direction of the cylindrical section (121) is 19-21 °.

7. A method for controlling the attitude of a launch vehicle according to any one of claims 1 to 6, comprising the steps of:

The 1 st to N-2 nd engines are utilized to push the carrier rocket to fly to a set height or to be separated after flying for a set time, and the fairing back cover (12) is separated;

The rest part of the fairing rear cover (12) is controlled to be 180 degrees reversed after being separated by using a gesture control mechanism;

the fairing front shroud (11) is thrown off, and the load (2) is continuously propelled by the N-1 stage engine (31) and the N stage engine (32).

8. A method of controlling the attitude of a launch vehicle according to claim 7, wherein the pitching maneuver or yawing maneuver is performed by the first attitude control assembly while controlling the attitude of the entire launch vehicle, and the rolling maneuver is performed by the second attitude control assembly with the second attitude control assembly as a backup.

9. A launch vehicle attitude control method according to claim 8, wherein, when performing the flight section attitude control before the rudder (6) is disengaged:

The control moment in the pitching direction is

Wherein l ₁ is the axial distance from the thrust line of the first engine block (41) to the arrow body centroid, F ₁ is the thrust of the first engine block (41), l ₂ is the axial distance from the thrust line of the second engine block (42) to the arrow body centroid, F ₂ is the thrust of the second engine block (42),Pitch control moment provided for the rudder; Is rudder deflection angle of The pneumatic pitching moment coefficient of the rocket is q is dynamic pressure, S is pneumatic reference area, L is pneumatic reference length,Is rudder deflection angle ofThe aerodynamic pitching moment coefficients of the rocket are u1 to u8, which are respectively on-off control instructions of the T1 to T8 engine units, wherein 1 is on or 0,1 is off, and 0 is on;

The control moment M _ψ of the yaw path is:

the control moment M _γ of the roll channel is:

When the attitude control of the flight section after the rudder (6) is disengaged is performed:

control moment in pitch direction The method comprises the following steps:

the control moment M _ψ of the yaw path is:

The control moment M _γ of the roll channel is:

10. A method of controlling the attitude of a launch vehicle according to claim 9, wherein said controlling the remaining 180 degrees of reversal of the cowling back (12) after detachment by the attitude control mechanism comprises: the remaining 180 degrees of the cowling back cover (12) is reversed after being separated by four first engine blocks (41) of the first attitude control assembly, and the attitude control and the backup are carried out by four second engine blocks (42) of the second attitude control assembly.