A two-stage reusable launch vehicle

Abstract

A two-stage reusable carrier comprises a first stage and a second stage connected in series; the first stage is an axisymmetric thin-walled cylindrical rocket body, and is provided with canards, lift wings and V-shaped tail wings; the lift wings are fixed wings, located at the rear of the rocket body, in the form of delta wings, and provide lift or resistance for the first stage during horizontal return, and elevator ailerons are arranged under the wing body, which can be used for steering or up and down yaw; the canards are single-axis swing wings, the rotation axis is parallel to the arrow axis, and are located at the front of the rocket body; the V-shaped tail wing is a fixed wing, located at the rear of the rocket body, and is in the form of symmetrical double tail wings; the second stage is an axisymmetric thin-walled cylindrical rocket body, and the layout scheme is similar to that of the first stage. The present invention provides a design scheme for the integration of traditional rockets and winged aircraft, which solves the problem that traditional rockets do not have the ability to fly horizontally on the one hand, and on the other hand, takes advantage of the high efficiency of the rocket configuration structure, and realizes the integration of the advantages of the two.

Description

Technical Field

[0001] The invention relates to a two-stage reusable carrier, belonging to the overall technical field of carrier rockets. Background Art

[0002] It is an inevitable trend for launch vehicles to move from disposable to reusable. Reusability must first solve the problem of lossless recovery of launch vehicle sub-stage modules. The essence of lossless recovery is to achieve energy dissipation during the return process under certain economic and capacity constraints. This energy dissipation method generally adopts aerodynamics, main power deceleration or a combination of them.

[0003] At present, the reuse of launch vehicles at home and abroad can only achieve the recovery and reuse of one sub-stage. The implementation method is generally to adopt the active power deceleration mode. In the landing process, the outrigger support method, the recovery net recovery method, and the hook recovery method can be adopted. The use of the engine power deceleration mode requires the launch vehicle to reserve a large amount of fuel in advance to provide power deceleration, which reduces the carrying efficiency. Moreover, this method is extremely inefficient in orbital stage recovery and cannot be used. How to effectively realize the recovery and reuse of two-stage rocket products, there is no mature technical solution in China. Summary of the invention

[0004] The technical problem to be solved by the present invention is: to overcome the shortcomings of the prior art and propose a two-stage reusable carrier connected in series, in which both the first and second stages adopt a winged fly-back method, which solves the problem of controllable recovery of the two-stage carrier and lays the foundation for the subsequent reuse of the two-stage carrier.

[0005] The purpose of the present invention is achieved through the following technical solutions:

[0006] A two-stage reusable carrier comprises a first sub-stage and a second sub-stage connected in series with each other;

[0007] The first stage is an axisymmetric thin-walled cylindrical body, and is equipped with canards, lift wings and V-shaped tail wings; the lift wings are fixed wings, located at the rear of the body, in the form of delta wings, providing lift or resistance for the first stage during horizontal return, and elevator ailerons are set under the wing, which can be used for steering or up and down yaw; the canards are single-axis swing wings, with the rotation axis parallel to the axis of the arrow, located at the front of the body; the V-shaped tail wing is a fixed wing, located at the rear of the arrow, in the form of symmetrical double tail wings;

[0008] The second stage is an axisymmetric thin-walled cylindrical rocket body, and its layout is similar to that of the first stage.

[0009] Preferably, the second stage is also provided with canards, lift wings and V-tail for aerodynamic control.

[0010] Preferably, a front landing gear is embedded in the front part of a sub-stage, folded in during ascent, and opened before landing; two sets of rear landing gear are embedded in the joint between the root of the lift wing and the rocket body, folded inside the lift wing and the rocket body during ascent, and opened before landing.

[0011] Preferably, an inflatable head assembly is provided at the front of a sub-stage, which is not inflated during ascent and is retracted inside the interstage section. It is inflated and expanded before the sub-stage re-enters the atmosphere to improve the aerodynamic characteristics.

[0012] Preferably, the front head of the second sub-stage is a fixed structure to form a rectifying effect.

[0013] Preferably, a front landing gear is embedded in the front part of the second stage, folded in during the ascent, and opened before landing; two sets of rear landing gear are embedded in the joint between the root of the lift wing and the rocket body, folded inside the lift wing and the rocket body during the ascent, and opened before landing.

[0014] Preferably, the middle front part of the second-stage fuselage serves as a load compartment, which can be used to adjust the center of mass position of the second-stage.

[0015] Preferably, both sub-stages use liquid oxygen-kerosene propellants; the first sub-stage separates from the second sub-stage after completing the ascent flight, and re-enters and recovers by means of a horizontal flyback; the second sub-stage re-enters and returns after completing the main mission of transporting the payload into orbit, and re-enters and recovers by means of a horizontal flyback.

[0016] Preferably, the first stage and the second stage take off vertically in series, the first stage and the second stage are separated, and the first stage enters the horizontal flight back stage, which is specifically divided into a high-altitude gliding attitude adjustment stage, a high-angle-of-attack reentry deceleration stage, a gliding stage before landing (and a landing stage; the second stage relies on the main power to work continuously, and after completing the predetermined payload release task, it completes the deceleration from the orbit under the action of its own auxiliary power, and reenters at a small inclination angle on the orbit, and after a long-distance high-altitude glide, it lands and returns near the designated landing site. The reentry flight profile is similar to the horizontal flight back of the first stage, and is divided into a high-angle-of-attack reentry deceleration stage, a terminal energy management stage, and an automatic landing stage.

[0017] Preferably, during the horizontal return process of a sub-stage, different control strategies are implemented according to different dynamic pressure conditions: in the initial re-entry phase, when the dynamic pressure is less than 100Pa, the auxiliary power system is used for attitude control, and as the dynamic pressure increases, the aerodynamic control surfaces begin to intervene; when the dynamic pressure further increases and the flight altitude decreases, the auxiliary power system finally exits attitude control, and attitude control is completed only by the aerodynamic control surfaces.

[0018] Compared with the prior art, the present invention has the following beneficial effects:

[0019] (1) The present invention provides a new two-stage tandem carrier configuration, in which both the first and second stages can adopt a winged configuration for horizontal return, thereby achieving recovery;

[0020] (2) The present invention provides a method for both the first and second stages to fly back horizontally, making full use of the effect of the atmosphere for energy dissipation, significantly reducing the difficulty of developing engines for vertical take-off and landing recovery. The vertical take-off and landing rockets currently under development all require deep throttling and multiple starts for the engines, which is difficult to tackle;

[0021] (3) The present invention provides an aerodynamic layout of canard + delta wing, which can effectively improve the aerodynamic control efficiency through the uniaxial swing of the canard;

[0022] (4) The present invention provides a design scheme that integrates traditional rockets and winged aircraft. On the one hand, it solves the problem that traditional rockets do not have the ability to fly horizontally. On the other hand, it takes advantage of the high efficiency of the rocket configuration structure, thus achieving a fusion of the advantages of the two.

[0023] (5) The present invention provides key design parameters for the reentry flight profiles of the first and second stages, and provides control schemes for different flight segments. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Figure 1 This is a configuration scheme of a two-stage reusable launch vehicle, including three views.

[0025] Figure 2 This is a schematic diagram of the control participation of additional aerodynamic control surfaces such as canards, lift wings and tail wings.

[0026] Figure 3 This is the cross-sectional parameter diagram of the first-stage flyback.

[0027] Figure 4 This is the cross-sectional parameter diagram of the secondary flyback.

[0028] Figure 5 This is the schematic diagram of the attitude control system.

[0029] Figure 6 This is the structural diagram of the composite attitude control system. DETAILED DESCRIPTION

[0030] To make the objectives, technical solutions and advantages of the present invention more clear, the embodiments of the present invention will be further described in detail below with reference to the accompanying drawings.

[0031] A two-stage reusable launch vehicle, including a first stage and a second stage connected in series; both stages use liquid oxygen-kerosene propellant, the first stage is equipped with 7 liquid oxygen-kerosene engines, and the second stage is equipped with 1 liquid oxygen-kerosene engine. The takeoff weight of the whole rocket is about 720 tons, and the takeoff thrust is about 8750kN. After the first stage completes the ascent flight, it separates from the second stage and reenters and recovers by horizontal return; after the second stage completes the main task of transporting the payload into orbit, it reenters and returns, and also reenters and recovers by horizontal return.

[0032] The first stage is an axisymmetric thin-walled cylindrical body, and is equipped with canards, lift wings and V-shaped tail wings; the lift wings are fixed wings, located at the rear of the body, in the form of large delta wings, providing lift or resistance for the first stage during horizontal return, and elevator ailerons are set under the wing, which can be used for steering or up and down yaw; the canards are single-axis swing wings, with the rotation axis parallel to the axis of the arrow, located at the front of the body; the V-shaped tail wing is a fixed wing, located at the rear of the body, in the form of symmetrical double tail wings. The front landing gear is embedded in the front of the first stage, folded during the ascent, and opened before landing; two sets of rear landing gear are embedded in the joint between the lift wing root and the body, folded inside the lift wing and the body during the ascent, and opened before landing. An inflatable head assembly is set at the front of the first stage, which is not inflated during the ascent and folded inside the interstage. It is inflated and unfolded before the first stage re-enters the atmosphere to improve the aerodynamic characteristics.

[0033] The second stage is an axisymmetric thin-walled cylindrical body with a similar layout to the first stage. It is equipped with canards, lift wings and V-shaped tail fins for aerodynamic control. The front landing gear is embedded in the front part, folded during ascent and opened before landing. Two sets of rear landing gear are embedded in the joint between the lift wing root and the body. They are folded inside the lift wing and the body during ascent and opened before landing. The difference from the first stage is that the front head of the second stage is a fixed structure, which creates a rectifying effect. The front center of the second stage fuselage serves as a load compartment, which can be used to adjust the center of mass position of the second stage.

[0034] The first and second stages take off vertically in series and fly in the mode of a conventional two-stage launch vehicle. When they reach an altitude of about 72km and a speed of about 2550m / s, the first stage separates from the second stage and enters the horizontal flight back phase, which is specifically divided into a high-altitude gliding attitude adjustment phase (about 125s), a high-angle-of-attack re-entry deceleration phase (about 73s), a pre-landing gliding phase (about 125s) and a landing phase. The peak overload of the aerodynamic deceleration phase does not exceed 9 gravitational accelerations. The second stage relies on the main power to continue working and enters a low-Earth orbit at an altitude of 200km. After completing the scheduled payload release mission, it can complete the deceleration from orbit under the action of its own auxiliary power, and re-enter the orbit with a small inclination angle. After a long distance of high-altitude gliding, it will land and return near the designated landing site. The re-entry flight profile is similar to the horizontal flight return of the first stage, which is divided into a high-angle re-entry deceleration section (starting from an altitude of about 120km to around 30km), a terminal energy management section (from around 30km to 3km), and an automatic landing section (from the end of the energy management section to a horizontal landing).

[0035] In the first stage ascent phase, the first stage main power engine swings to control pitch, yaw and roll. In the second stage ascent phase, the second stage main power is used for pitch and yaw control, and auxiliary power is used for roll control. In the first stage horizontal return process, different control strategies are implemented according to different dynamic pressure conditions. In the initial reentry phase, when the dynamic pressure is less than 100Pa, the auxiliary power system is used for attitude control. As the dynamic pressure increases, the aerodynamic control surfaces begin to intervene. When the dynamic pressure increases further and the flight altitude decreases, the auxiliary power system finally exits attitude control, and attitude control is completed only by the aerodynamic control surfaces.

[0036] Example:

[0037] 1) The two-stage reusable launch vehicle configuration scheme is shown in Figure 1 , as follows:

[0038] 11) The carrier configuration scheme adopts a two-stage tandem launcher, both of which use liquid oxygen-kerosene propellants. The first stage is equipped with 7 liquid oxygen-kerosene engines, with a single thrust of 1250kN, and the second stage is equipped with 1 liquid oxygen-kerosene engine. The takeoff weight of the whole rocket is about 720 tons, and the takeoff thrust is about 8750kN. The main body of the first stage is an axisymmetric thin-walled cylindrical rocket body with a diameter of 5m and a length of about 39m. It is equipped with a liquid oxygen tank and a kerosene tank. The length of the liquid oxygen tank is about 19m, and the length of the kerosene tank is about 11m; the main body of the second stage is an axisymmetric thin-walled cylindrical rocket body with a diameter of 5m and a length of about 33m. It is equipped with a liquid oxygen tank and a kerosene tank. The length of the liquid oxygen tank is about 7m, and the length of the kerosene tank is about 3.7m. The front half of the second stage is equipped with a payload cabin, with a cabin length of about 13m and a usable diameter of about 4.2m.

[0039] 12) The first stage adds canards, lift wings and V-shaped tail wings to the traditional rocket configuration. The lift wings are fixed wings located at the rear of the rocket body, with a symmetrical layout. There are two of them in the form of large delta wings with an area of ​​no less than 70m 2 The delta wing has a wingspan of 9.7m, a width of 12m near the rocket root, a wingtip width of 2.8m, and a Naka airfoil cross-section. There are two elevator ailerons under the lift wing, with an area of ​​no less than 6m 2 The canard is located at the front of the first stage, with a symmetrical layout. There are two pieces, and the leading edge is flush with the first and second stage separation surface. The canard has a wingspan of 3.3m, a width of 9m near the root of the rocket, and a wing tip width of 2m, with an area of ​​no less than 18m. 2 The canard can rotate around the X-axis of the arrow system and is controlled when folded and opened. The V-shaped tail is located at the tail, with a symmetrical layout. There are two of them, and the area of ​​a single tail is about 15m 2 The schematic diagram of the control of aerodynamic surfaces such as canards, lift wings and tail wings is shown in Figure 2 .

[0040] 13) The second stage cancels the traditional rocket fairing, integrates the payload compartment with the traditional rocket body tank, and adds canards, lift wings and V-shaped tail wings on its basis. The overall structural layout is similar to the first stage, only the corresponding wing area is adaptively adjusted. The lift wing area is about 35m 2 There are two elevator ailerons under the lift wing, with an area of ​​not less than 3m 2 The canard area is not less than 10m 2 The canard can rotate around the X-axis of the arrow system and is controlled when folded and opened. The V-shaped tail is located at the tail, with a symmetrical layout. There are two of them, and the area of ​​a single tail is about 8m 2 .

[0041] 14) Both the first and second stages are equipped with landing gears. The landing gear adopts the mature aircraft landing gear system solution. The first stage is equipped with three sets, one set of front landing gear is set in the first interstage section, and two sets of rear landing gear are set at the junction of the lift wing and the rocket body. They are folded during the ascent and return process, and opened before landing. The second stage is also equipped with three sets, located at the front and rear.

[0042] 15) An interstage is provided at the separation surface between the first and second stages, and the second stage engine is wrapped and protected inside the interstage. At the same time, an inflatable head assembly is provided in the interstage of the first stage. It is not inflated during the ascent process and is retracted inside the interstage. It is inflated and expanded before the first stage re-enters the atmosphere to improve the aerodynamic characteristics of the first stage re-entry.

[0043] 2) The flight profile of the two-stage reusable launch vehicle is as follows:

[0044] The ascent process adopts a vertical takeoff mode, and goes through a first-stage flight phase and a second-stage flight phase. Its basic process is consistent with that of a traditional two-stage launch vehicle. For the return process of the first and second stages, a detailed design of the flight profile is carried out. The return profile of the first stage adopts the idea of ​​"high-altitude rapid energy consumption + low-altitude balanced gliding" to manage the energy of the re-entry phase and optimize the environmental conditions at the time of landing.

[0045] 21) The flight profile of the first-stage return flight is divided into the high-altitude gliding attitude adjustment stage, the high-angle-of-attack reentry deceleration stage, the pre-landing gliding stage and the landing stage. The detailed flight profile curve is shown in Figure 3 The main task of the high-altitude taxiing attitude adjustment stage is to provide a better initial attitude for entering the next stage, while ensuring that the initial dynamic pressure, peak heat flow, etc. are within a reasonable range; the high-angle-of-attack reentry deceleration mainly uses aerodynamics to decelerate, and achieves a significant reduction in speed under the constraints of overload and heat flow. Peak overload and peak heat flow may occur during the deceleration process; the pre-landing gliding stage completes two tasks, one is to complete the low-altitude pull-up, and the other is to continuously reduce the altitude by a small amount with a relatively certain attitude and speed inclination, providing better conditions for the final safe landing; the landing stage opens the landing gear for landing.

[0046] High-altitude gliding attitude adjustment phase: The separation point of the first and second stages is generally 72km, with a speed of about 2550m / s and a local ballistic inclination of about 21°. After the first stage separates, it enters the high-altitude gliding attitude adjustment phase. The atmosphere is thin during this phase, and aerodynamic effects are basically not considered. The aircraft flies in a parabola, and the attitude is adjusted by the auxiliary power system during gliding to ensure that the attitude meets the conditions when re-entering the atmosphere;

[0047] High-angle-of-attack reentry deceleration phase: After about 125 seconds of gliding flight, the first stage enters the high-angle-of-attack reentry deceleration phase. The reentry conditions are generally 50km altitude, about 2580m / s speed, and the local ballistic inclination is about -24°. In the early stage of reentry, the speed does not change much, the dynamic pressure is very small, and the heat flow has the greatest impact on it. Therefore, in order to slow down the excessive heat flow caused by the rapid descent in altitude, a high-angle-of-attack flight is used to generate greater lift and drag, while effectively decelerating. In this phase, the maximum angle of attack α is usually used first. max As the flight altitude decreases, the atmospheric density increases. At the end of the reentry phase, the optimal lift-to-drag ratio angle of attack α is adopted. LDmax Flying can effectively increase the range. Therefore, the attack angle profile of the reentry vehicle is determined to be in the form of segmented constants, and linearly changing attack angles are generally used between segments to achieve smooth connection. The attack angle profile has the following form, as shown in the following formula.

[0048]

[0049] Where V 1 、V 2 is a constant value, see Figure 3In this scheme, the maximum reentry angle of attack is set to 42°, V 1 The value is 2500m / s, V 2 The value is 0.7 Mach, the optimal lift-to-drag ratio angle of attack α L / Dmax The end point of the high-angle-of-attack reentry deceleration phase is about 15 km high, with a speed of about 280 m / s and a local trajectory inclination of about -27.5°.

[0050] Quasi-balanced gliding phase before landing: When entering the quasi-balanced gliding phase, this patent adopts a trajectory design method based on the angle of attack-pseudo-dynamic pressure profile. The general principle of this method is as follows:

[0051]

[0052] E is the total energy of the aircraft, m is the total mass of a sub-stage, V is the flight speed, g is the acceleration of gravity, and h is the flight altitude. As shown in the above formula, the representation of the total flight energy is closely related to the flight speed V and the flight altitude h. In addition, the dynamic pressure q can also be represented as a function of V and h. Based on the level flight assumption, the dynamic pressure in the balanced gliding flight phase can actually be expressed as a function of the altitude h. By converting the "angle of attack-dynamic pressure" profile into the "angle of attack-altitude (pseudo-dynamic pressure)" profile, an angle of attack design method is formed.

[0053]

[0054] In the above formula, α is the angle of attack, h is the flight altitude, and A and B are related parameter values. Based on the quasi-balanced glide angle design method, the final angle of attack changes from 5° to about 17°, and the flight speed decreases from 280m / s to 70m / s.

[0055] Landing stage: At an altitude of 3km from the ground, open the landing gear and land. The landing conditions must not exceed the landing speed of a civil airliner.

[0056] 22) Second level return flight profile

[0057] The second stage reenters from orbit, which is characterized by a very high speed (about 7500m / s) and a large vertical acceleration. As the altitude decreases, the air density increases, the dynamic pressure increases exponentially, and the overload also increases accordingly. The dynamic pressure and overload can easily exceed the physical strength limit. Especially the dynamic pressure. Since the sinking rate during the return process is large, the aerodynamic force is unable to directly control the dynamic pressure, so it is necessary to maintain a large angle of attack from the beginning of the return (to suppress the trend of increasing dynamic pressure). At the same time, a large angle of attack descent will cause the overload to exceed the constraint first. To avoid structural damage, a fixed overload flight mode is adopted when the overload reaches a predetermined value. As long as the fixed overload meets the overall design constraints, the overload constraint problem can be solved. On this basis, through the effect of large overload, the trajectory will gradually pull up, the sinking rate will be reduced, and the dynamic pressure constraint will also be met. See the detailed flight profile curve. Figure 4 .

[0058] Angle of attack holding phase: The angle of attack holding phase starts from returning to the highest point. The second stage first establishes a high angle of attack flight state and descends at a fixed constant angle of attack. As the altitude decreases, the dynamic pressure increases, and the overload increases accordingly. In this scheme, the second stage adopts a 40° high angle of attack for about 1800s, reducing the speed from 7500m / s to about 280m / s, and the altitude from 120km to about 40km.

[0059] Overload holding stage: When the overload reaches the preset value, it enters the overload holding stage. In the overload holding stage, the flight is kept at a fixed overload. As the dynamic pressure increases further, the flight attack angle gradually decreases, and the flight trajectory angle also gradually decreases.

[0060] Attack angle transition section: When the sinking rate reaches the predetermined value, it enters the attack angle transition section. At this time, the attack angle is still large and is behind the maximum lift-to-drag ratio. In the attack angle transition section, adjust the attack angle and overload to make the attack angle in front of the maximum lift-to-drag ratio.

[0061] Quasi-balanced glide stage: When the angle of attack is small and in front of the maximum lift-to-drag ratio, it enters the quasi-balanced glide stage. The quasi-balanced glide stage means maintaining the overload at 1g, using aerodynamic drag to further consume energy and ensure entering the energy management window (altitude 25km);

[0062] Landing stage: At an altitude of 3km from the ground, open the landing gear and land. The landing conditions must not exceed the landing speed of a civil airliner.

[0063] 3) The control scheme during flight is as follows:

[0064] The configuration scheme of the attitude control system during the flight of the first and second stages is basically the same, and both adopt the control scheme of three inertial groups + control combination + servo + attitude control nozzle + air rudder. A three-inertial group + altimeter radar + satellite composite guidance system is adopted. Before launching, the initial attitude reference is established through the self-aiming of the inertial group for inertial navigation. The radar altimeter works in the final stage of gliding. When it is a certain distance away from the target point, it performs altimeter / terrain matching. After completing the navigation error correction, it presses down to attack, thereby ensuring the accuracy of the landing point; at the same time, the combined navigation correction is started under the condition that the satellite is available to further improve the accuracy. The control system uses the three-inertial group to measure the attitude of the projectile, and the flight angle of attack, sideslip angle and roll angle are obtained through real-time calculation. After the control law is solved and the control signal is output, the attitude control nozzle switch and the air rudder deflection are controlled to generate control torque, thereby ensuring attitude stability and tracking the guidance command. The control system scheme is shown in Figure 5 .

[0065] 31) The return process is a three-channel decoupling control, which adopts the attitude angle deviation, angular velocity, variable gain, and variable network three-axis stable control method. The three-channel control equation is:

[0066]

[0067]

[0068]

[0069] in: Static gains for pitch, yaw and roll channels;

[0070] discrete correction networks for pitch, yaw, and roll channels;

[0071] W GZ (s),W ST (s),W JB (s) is the dynamic characteristics of the strapdown inertial measurement unit, rate gyro, and strapdown meter;

[0072] Δα 1 ,β,Δγ v are the angle of attack, sideslip angle and velocity roll angle deviation in the rocket body coordinate system;

[0073] is the dynamic gain of the pitch, yaw and roll channels, which is the ratio of the static gain of the attitude angular velocity channel to the static gain of the attitude angle channel;

[0074] ω z1STR ,ω y1STR ,ω x1STR are the pitch, yaw and roll angular velocities about the rocket body coordinate system.

[0075] The attitude control system adopts the BTT control method in which the angle of attack and roll angle track the guidance instructions and the guidance sideslip angle instruction is zero, and the auxiliary power / air rudder composite control strategy.

[0076] Auxiliary power control: The pitch channel adopts the angle of attack deviation control scheme, the yaw channel adopts the sideslip angle feedback control scheme, and the roll channel adopts the roll angle deviation control scheme.

[0077] Air rudder control: The pitch channel adopts the attack angle deviation + pitch angular velocity control scheme, the yaw channel adopts the sideslip angle + yaw angular velocity control scheme, and the roll channel adopts the roll angle deviation + roll angular velocity control scheme.

[0078] As the aircraft's altitude decreases and the dynamic pressure increases, the auxiliary power / air rudder composite control strategy transitions from three channels completely controlled by auxiliary power to air rudder main control and auxiliary power auxiliary control, and finally three channels are completely controlled by air rudder.

[0079] 32) Composite control strategy

[0080] The auxiliary power / pneumatic rudder composite control system consists of three parts: attitude controller, composite control strategy, and auxiliary power and pneumatic rudder control distribution. Figure 6 As shown in the figure, the attitude controller obtains the control torque command M based on the attitude angle command and the sensor's measured attitude angle, attitude angular rate and other signals. c The composite control strategy distributes the torque instructions of the auxiliary power and the pneumatic rudder according to the current dynamic pressure, the working conditions of the thruster and the pneumatic rudder. The control distribution algorithm of the auxiliary power and the pneumatic rudder realizes the mapping of the control torque instruction into the specific thruster switch instruction and the deflection instruction of the rudder. The aircraft drives the actuator to work according to the instruction of the actuator, and then controls the aircraft to achieve the required flight state.

[0081] 33) Segment control mode

[0082] According to the dynamic pressure and Mach number, the entire reentry flight process is divided into five control modes. The specific reentry control modes are shown in the following table.

[0083] Table 1

[0084]

[0085]

[0086] Where: T R,r ,T R,p ,T R,y RCS thrusters for roll, pitch, and yaw channels; a ,δ e ,δ r ,δ f They are the equivalent aileron, equivalent elevator, rudder and canard deflection surfaces respectively.

Claims

1. A two-stage reusable carrier, characterized in that: It includes a first sub-stage and a second sub-stage connected in series with each other; The first stage is an axisymmetric thin-walled cylindrical body, and is equipped with canards, lift wings and V-shaped tail wings; the lift wings are fixed wings, located at the rear of the body, in the form of delta wings, providing lift or resistance for the first stage during horizontal return, and elevator ailerons are set under the wing, which can be used for steering or up and down yaw; the canards are single-axis swing wings, with the rotation axis parallel to the axis of the arrow, located at the front of the body; the V-shaped tail wing is a fixed wing, located at the rear of the arrow, in the form of symmetrical double tail wings; The second stage is an axisymmetric thin-walled cylindrical body, and is also equipped with canards, lift wings and V-shaped tail fins for aerodynamic control. The front landing gear is embedded in the front of the first stage, which is folded during the ascent and opened before landing. Two sets of rear landing gear are embedded in the joint between the lift wing root and the rocket body, which are folded inside the lift wing and the rocket body during the ascent and opened before landing. An inflatable head assembly is provided at the front of the first stage. It is not inflated during the ascent and is retracted inside the interstage. It is inflated and deployed before the first stage re-enters the atmosphere to improve the aerodynamic characteristics.

2. The two-stage reusable vehicle according to claim 1, characterized in that: The front head of the second stage is a fixed structure, which forms a rectifying effect.

3. The two-stage reusable vehicle according to claim 1, characterized in that: The front landing gear is embedded in the front part of the second stage, which is folded in during the ascent and opened before landing. Two sets of rear landing gear are embedded in the junction of the lift wing root and the rocket body, which are folded inside the lift wing and the rocket body during the ascent and opened before landing.

4. The two-stage reusable vehicle according to claim 1, characterized in that: The middle front part of the second stage fuselage serves as a payload bay, which can be used to adjust the center of mass position of the second stage.

5. The two-stage reusable vehicle according to any one of claims 1 to 4, characterized in that: Both sub-stages use liquid oxygen and kerosene propellants; the first sub-stage separates from the second sub-stage after completing the ascent flight and re-enters and recovers by horizontal flyback; the second sub-stage re-enters and returns after completing the main mission of transporting the payload into orbit, and re-enters and recovers by horizontal flyback.

6. The two-stage reusable vehicle according to any one of claims 1 to 4, characterized in that: The first stage and the second stage take off vertically in series, separate from each other, and the first stage enters the horizontal flight back stage, which is specifically divided into the high-altitude gliding attitude adjustment stage, the high-angle re-entry deceleration stage, the pre-landing gliding stage and the landing stage; the second stage relies on the main power to work continuously, and after completing the predetermined payload release task, it completes the deceleration from orbit under the action of its own auxiliary power, and re-enters the orbit at a small inclination angle. After a long-distance high-altitude gliding, it lands and returns near the designated landing site. The re-entry flight profile is divided into the high-angle re-entry deceleration stage, the terminal energy management stage, and the automatic landing stage.

7. The two-stage reusable vehicle according to claim 6, characterized in that: During the horizontal return process of the first stage, different control strategies are implemented according to different dynamic pressure conditions: in the initial re-entry phase, when the dynamic pressure is less than 100Pa, the auxiliary power system is used for attitude control. As the dynamic pressure increases, the aerodynamic control surfaces begin to intervene; when the dynamic pressure increases further and the flight altitude decreases, the auxiliary power system finally exits attitude control, and attitude control is completed only by the aerodynamic control surfaces.

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