CN118323477A - Reusable horizontal take-off and landing two-stage aerospace carrier - Google Patents

Abstract

The invention provides a reusable horizontal take-off and landing two-stage aerospace vehicle, wherein the primary structure is a reusable winged sub-orbit aircraft, and the whole aerospace vehicle is in an aerospace plane configuration; the secondary structure is a traditional rocket configuration. Wherein the first stage can realize horizontal lifting, and the second stage adopts a recovery scheme of offshore splash recovery. The invention realizes the recovery and the reusability of the aircraft while ensuring the high-efficiency carrying.

Description

Reusable horizontal take-off and landing two-stage aerospace carrier

Technical Field

The invention belongs to the field of aerospace, and particularly relates to a reusable horizontal take-off and landing two-stage aerospace carrier.

Background

An aerospace vehicle (aerospace vehicles, ASV) refers to a novel aircraft capable of realizing aerospace navigation, integrates the technology and characteristics of an aerospace vehicle and an aerospace vehicle, can well meet the requirements of frequent, reliable and economic aerospace transportation in the future, and can fly in the near space which is difficult to reach by a general aircraft.

Compared with the traditional rocket-propelled spacecraft, the aerospace craft has the following advantages:

1. The method has the advantages that a vertical emission base is not needed, the airport is utilized to realize horizontal take-off and landing, the emission preparation time is greatly reduced, the method has the characteristics of short emission period and high reaction speed, and the method can realize future high-frequency and large-scale flight space transportation and can rapidly execute emergency tasks;

2. The rocket-based combined cycle or precooling air suction type cycle and other propulsion modes are adopted, so that the rocket-based combined cycle or precooling air suction type cycle has good specific impulse characteristics and economic characteristics, and is beneficial to improving the carrying efficiency;

3. meanwhile, the device has reusability, adopts single-stage or two-stage rail feeding, and can greatly reduce the emission cost;

4. the flying speed reaches hypersonic speed (Ma is more than 5), and the flying envelope is wide, and has strong burst prevention capability and survivability.

Aerospace vehicles offer great convenience for space delivery, but at the same time present a number of challenges. For example, an aerospace vehicle flies at hypersonic speed, and needs to reenter the atmosphere, and meanwhile, the aerospace vehicle has high reusability, so that a great challenge is faced in terms of heat protection; at present, an aerospace vehicle is limited by specific impact of an engine and structural quality, single-stage rail entering is difficult to realize, a trajectory scheme is required to be accurately designed and optimized by adopting two-stage rail entering, and meanwhile, the problems of horizontal take-off and landing of one stage, recovery of two stages and the like are required to be solved.

Disclosure of Invention

The invention provides a reusable horizontal take-off and landing two-stage aerospace vehicle, which is provided with a primary structure of a reusable winged sub-orbit aircraft and is integrally in an aerospace plane configuration, and aims to solve the technical problems by researching the current advanced aerospace carrying scheme from home and abroad and facing the application background of the aerospace shuttle of the cross-domain flight from the ground to the near-ground orbit; the secondary structure is a traditional rocket configuration. The first stage can realize horizontal take-off and landing, the second stage adopts a recovery scheme of marine splash recovery, and the recovery and the repeated use of the aircraft are realized while the high-efficiency carrying is ensured. Meanwhile, in order to verify the rationality and feasibility of the scheme, the result is analyzed and discussed through a ballistic simulation and optimization model.

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

A reusable horizontal take-off and landing two-stage space carrier comprises a primary structure and a secondary structure; the primary structure is an aerospace plane configuration and is used for flushing out the atmosphere and sending the effective load and the secondary structure into the sub-track; the secondary structure is a rocket configuration without lifting surfaces and is symmetrically distributed on the whole and is used for conveying effective loads into a preset track; the primary structure comprises a machine head, a load cabin, an air inlet channel and a primary engine; the load cabin is positioned at the back of the middle section of the machine head and is used for accommodating the secondary structure and the effective load; the air inlet channel is positioned at the front part of the belly and the wing and is used for decelerating air flow and providing air inlet for the primary engine; the primary engine is positioned at the rear part of the belly of the rear section of the aircraft nose and behind the air inlet channel and provides thrust for the power flight section with a primary structure; the secondary structure comprises a secondary main engine, an inflatable heat shield and a recovery parachute, wherein the recovery parachute is used for heat protection when the secondary structure reenters and returns and providing buoyancy, and the recovery parachute is used for final-stage deceleration of the secondary structure.

Further, the primary structure also comprises wings, movable duckwings, vertical tail wings, a front liquid hydrogen storage tank, a rear liquid hydrogen storage tank and a primary liquid oxygen storage tank; the movable duck wing is positioned on two sides of the aircraft nose and at the front part of the aircraft wing after the middle parts of the two sides of the aircraft nose are deviated, the vertical tail wing is positioned at the tail end of the aircraft wing, the air inlet channel is positioned on the belly and the front part of the aircraft wing, the primary liquid oxygen storage tank is positioned on two sides of the load cabin, the front liquid hydrogen storage tank is positioned in the front cone of the aircraft nose, and the rear liquid hydrogen storage tank is positioned behind the load cabin.

Further, the secondary structure also comprises a secondary liquid oxygen storage tank, a liquid hydrogen storage tank, a tank interval, a payload adapter ring, a posture rail control engine and an IVF system; the liquid hydrogen storage tank is arranged at the front, the secondary liquid oxygen storage tank is arranged at the rear, and the liquid hydrogen storage tank is connected with the secondary liquid oxygen storage tank through a tank interval to provide propellant for the secondary main engine; the secondary main engine is arranged in the middle of the secondary liquid oxygen storage tank, the effective load connecting ring is connected with the effective load and the liquid hydrogen storage tank, the attitude and orbit control engine is positioned at the rear part of the liquid oxygen storage tank and used for attitude and orbit control of a secondary structure, the inflatable heat shield is positioned at the inner side of the effective load connecting ring and between the effective load and the liquid hydrogen storage tank, and the recovery parachute is positioned at the rear part of the secondary liquid oxygen storage tank.

Further, an adjustable inclined plate is arranged on the inner side of the air inlet channel and is adjusted according to the flight Mach number, so that the air flow meets the air inlet requirement of the primary engine; when the first-stage engine works in a rocket mode, the air inlet is completely closed by the adjustable inclined plate of the air inlet.

Furthermore, the primary engine tightly combines three power cycles of a turbine, a ram and a rocket together to realize the work in full airspace and full speed.

Further, the landing gear further comprises a front landing gear, a main landing gear and a rear landing gear, and the landing gear is used for supporting the horizontal take-off and landing two-stage space carrier on the runway.

Further, the secondary liquid oxygen storage tank is annular.

Further, the IVF system provides power for the secondary structure, provides gaseous propellant for the attitude and orbit control engine, provides high-pressure gas for the inflatable heat shield, and manages the propellant in each storage tank.

Further, after the primary structure is ignited, the runway is horizontally taken off; retracting the movable duck wings at 0.9Ma to reduce drag; when reaching 6Ma, switching the first-stage engine from an air suction mode to a rocket mode; when the height reaches 90km, the primary structure is shut down, and the secondary structure and the effective load are discharged from the load cabin; the secondary main engine is then ignited, delivering the payload into the predetermined track.

Further, after the primary structure and the secondary structure are separated, the primary structure enters the atmosphere after transient ballistic sliding, is decelerated through gliding, and falls off to return to a runway in take-off; the secondary structure is separated from the effective load, the effective load is discarded, the effective load enters a return trajectory through the braking of a secondary main engine, then an inflatable heat shield is unfolded and enters an atmosphere, a recovery parachute at the tail part is unfolded after full deceleration, finally the effective load is splashed on the sea surface, the buoyancy provided by the inflatable heat shield floats on the sea surface, and the effective load is recovered.

Compared with the prior art, the invention has at least one of the following beneficial effects:

1. the primary structure and the secondary structure can be repeatedly used for many times;

2. The carrying coefficient of the aircraft is high; the carrying capacity of the carrier 200kmLEO is 12t, and the load coefficient is up to 5 percent, which is higher than 4.15 percent of falcon No. 9 and 3.2 percent of the load coefficient of a single-stage in-orbit carrier in China;

3. the aircraft has high flexibility, and the primary structure and the secondary structure can be used independently;

4. The liquid hydrogen and liquid oxygen are adopted as the propellant, so that the environment is protected.

Drawings

FIG. 1a is a flow chart of a design of a primary structure of a reusable level take-off and landing two-stage aerospace vehicle of the present invention;

FIG. 1b is a flow chart of a design of a secondary structure of a reusable level lift secondary space carrier of the present invention;

FIG. 2 is a schematic diagram of a primary structure;

FIG. 3 is a schematic diagram of a secondary structure;

FIG. 4 is a specific impulse graph of an engine;

FIG. 5a, FIG. 5b illustrates the trajectory and parameters of the flight at the working stage; wherein, fig. 5a is a launching trajectory of a 200kmLEO track, and fig. 5b is a graph of the change of the flight speed and the acceleration of the working section with time;

FIG. 6a, FIG. 6b is a first order return trajectory and parameters; wherein, FIG. 6a is a first order return trajectory, and FIG. 6b is a first order return velocity and acceleration versus time plot;

FIG. 7a, FIG. 7b is a two-level return trajectory and parameters; wherein, fig. 7a is a secondary return trajectory, and fig. 7b is a secondary return velocity and acceleration versus time graph.

Wherein, the reference numerals are as follows: 1-payload, 2-payload bay, 3-tank bay, 4-secondary liquid oxygen tank, 5-secondary main engine, 6-IVF system, 7-liquid hydrogen tank, 8-recovery parachute, 9-nose, 10-duck wing, 11-load hatch, 12-vertical tail, 13-primary engine, 14-rear landing gear, 15-main landing gear, 16-wing, 17-air intake, 18-nose landing gear.

Detailed Description

The present invention will be described in further detail with reference to the drawings and examples, in order to make the objects, technical solutions and advantages of the present invention more apparent. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the scope of the invention. In addition, the technical features of the embodiments of the present invention described below may be combined with each other as long as they do not collide with each other.

As shown in fig. 1a and fig. 1b, the design concept of the reusable horizontal take-off and landing two-stage aerospace vehicle of the invention is as follows: firstly, determining a task target; secondly, determining basic performance and general configuration of an engine, developing detailed design on the appearance and structure of the carrier, performing fluid mechanical simulation (CFD) on the aerodynamic appearance to obtain aerodynamic characteristics, and performing ballistic simulation by using the aerodynamic characteristics; and finally, analyzing whether the ballistic simulation result can meet the task requirement. If not, the appearance and the structure are modified, and the subsequent flow is repeated until the requirements are met, so that the design is completed.

In the present invention, the design of the secondary structure is subordinate to the design of the primary structure. When designing the secondary structure, determining the theoretically required speed increment of the secondary structure according to the ballistic simulation result of the primary structureAnd a substantially global mass distribution relationship. And then finishing the model selection of the engine with the secondary structure, developing structural design and ballistic simulation, and performing iterative optimization to meet the requirements.

Specifically, taking a high-performance aerospace vehicle with the transport capacity not lower than 10 tons as an example, as shown in fig. 2, the reusable horizontal take-off and landing two-stage aerospace vehicle provided by the invention consists of a two-stage structure, and the used propellants are liquid hydrogen and liquid oxygen.

The primary structure is an aerospace plane configuration, and the main task is to flush out the atmosphere and send the payload 1 and the secondary structure into a sub-orbit. The primary structure mainly comprises a nose 9, a duck wing 10, a load cabin cover 11, a vertical tail wing 12, a primary engine 13, a rear landing gear 14, a main landing gear 15, a wing 16, an air inlet channel 17 and a nose landing gear 18.

The overall aerodynamic configuration employs a tailless delta wing with movable duck wings 10.

Preferably, the wing 16 is located at the middle of two sides of the nose 9 and is responsible for providing most of the lift force, and the wing profile adopts NACA0403 as a delta wing.

Preferably, the movable said duck wings 10 are located on both sides of the nose 9, in front of the wing 16. The movable duck wings 10 extend at low speed to provide additional lift and retract into the fuselage at high speed to reduce drag.

Preferably, the vertical tail 12 is located at the end of the wing 16 and is divided into two pieces, an upper piece and a lower piece, to provide course stability for flight in the atmosphere. The vertical tail 12 at the end of the wing 16 may reduce the induced drag while reducing the overall height of the vehicle.

Preferably, the inlet duct 17 is located in the belly and is provided in the front of the wing 16 for decelerating the airflow and providing inlet air to the primary engine 13. The inner side of the air inlet channel 17 is provided with an adjustable inclined plate which can be adjusted according to the flight Mach number so that the air flow meets the air inlet requirement of the engine. At supersonic speeds, the shock wave generated by the air intake 17 may provide additional lift. When the primary engine 13 operates in rocket mode, the adjustable inclined plate of the air inlet 17 can completely close the air inlet 17.

Preferably, the main body of the handpiece 9 is cylindrical and is configured to accommodate an internal structure comprising: a front liquid hydrogen storage tank, a rear liquid hydrogen storage tank, a load compartment, an engine compartment, a primary engine 13, a nose landing gear 18, a main landing gear 15, and a rear landing gear 14. The front part of the nose 9 is designed as an asymmetric cone, which can provide additional lift during supersonic flight.

Preferably, the load compartment is located at the back of the middle section of the nose 9, and accommodates the secondary structure and the payload 1, and the load compartment is provided with an upwardly opened load compartment cover 11, so that the secondary structure and the payload 1 can be released to the external environment according to task requirements.

Preferably, the two sides of the load compartment are provided with a primary liquid oxygen storage tank for containing liquid oxygen required by the operation of the primary engine 13.

Preferably, the front liquid hydrogen storage tank is positioned in the front cone of the machine head 9, and the rear liquid hydrogen storage tank is positioned behind the load compartment, and the front liquid hydrogen storage tank and the rear liquid hydrogen storage tank are used for containing liquid hydrogen required by the operation of the primary engine 13.

Preferably, the primary engine 13 is located at the rear part of the belly of the nose 9 and behind the air inlet channel 17, and provides thrust for the power flight section with a primary structure. The primary engine 13 tightly combines three power cycles of a turbine, a ram and a rocket together, and can realize the full-space-domain and full-speed-domain operation.

The structure of the primary engine 13 is divided into a precooler, a core machine and a ram outer duct. In the air suction mode, an air suction working mode of 'a turbine is adopted as a main part and punching is adopted as an auxiliary part', high-temperature air entering through a variable air inlet channel 17 is divided into two paths, one path of air enters into a pre-combustion chamber and a main combustion chamber in sequence under the action of a gas compressor after being cooled by a precooler, and is ejected at a high speed through a rocket nozzle to generate main thrust; the other path directly enters the punching outer duct to burn with a small amount of hydrogen, so as to generate partial thrust and improve the utilization rate of the hydrogen. Compared with the common turbine and ramjet combined engine, the thrust-weight ratio has obvious advantages under the condition that the specific impulse is basically equivalent.

When flying in rocket mode, the primary engine 13 adopts oxyhydrogen rocket mode to accelerate the aircraft to the orbit speed. The two working modes of the primary engine 13 share key components such as a turbine pump, an adjustable tail nozzle and the like, so that the structural quality of the engine is effectively reduced.

Preferably, the nose landing gear 18, the main landing gear 15, and the rear landing gear 14 are used for supporting a horizontal take-off and landing two-stage space carrier on a runway.

As shown in fig. 2 and 3, the secondary structure of the horizontal take-off and landing secondary space carrier is a conventional rocket configuration with no lifting surface and symmetrical overall layout, and is responsible for feeding the payload 1 into a predetermined orbit. The secondary structure mainly comprises a secondary liquid oxygen storage tank 4, a liquid hydrogen storage tank 7, a tank interval 3, a secondary main engine 5, a payload adapter ring 2, an attitude and orbit control engine, an IVF (INTEGRATED VEHICLE Fluids on an arrow) system 6, an inflatable heat shield and a recovery parachute 8.

Preferably, the liquid hydrogen tank 7 is preceded and the secondary liquid oxygen tank 4 is followed by a tank compartment 3 for supplying the secondary main engine 5 with propellant. The secondary liquid oxygen storage tank 4 is annular, and the secondary main engine 5 is arranged in the middle of the secondary liquid oxygen storage tank 4, so that the length of the secondary structure can be shortened.

Preferably, the payload collar 2 connects the payload 1 to a liquid hydrogen reservoir 7.

Preferably, the attitude and orbit control engine is used for attitude and orbit control of a secondary structure.

Preferably, the IVF system 6 serves a number of functions including providing electrical power to the secondary structure, providing gaseous propellant to the attitude and orbit engine, providing high pressure gas to the inflatable heat shield and managing the propellant in each tank.

Preferably, the inflatable heat shield is positioned inside the payload collar 2 between the payload 1 and the liquid hydrogen tank 7 for thermal protection during reentry and return of the secondary structure and to provide buoyancy.

Preferably, the recovery parachute 8 is positioned at the rear part of the secondary liquid oxygen storage tank 4 and used for decelerating the tail section of the secondary structure so that the recovery parachute can safely splash on the sea surface.

The working process of the reusable horizontal take-off and landing two-stage aerospace carrier comprises the following steps of:

The primary ignition flow is shown in table 1, and takes off horizontally from the runway; retracting the movable duck wings 10 at 0.9Ma to reduce drag; when reaching 6Ma, the first-stage engine 13 is switched from an air suction mode to a rocket mode; when reaching 90km, the primary shutdown is performed, and the secondary structure and the effective load 1 are discharged from the load cabin. The secondary main engine 5 is then ignited, feeding the payload 1 into the predetermined track.

TABLE 1

1. After the secondary separation, the primary structure enters the atmosphere after short ballistic sliding, is decelerated by gliding, and falls off the head to return to the runway during take-off. The secondary emission flow is shown in table 2, and the emission process is as follows: the bottoming engine ignites to bottom the propellant, and then the secondary main engine 5 ignites to accelerate the derailment. After entering the predetermined orbit, the payload 1 is separated first, and after the payload 1 is far away, the payload collar 2 is separated and the inflatable heat shield is deployed. When the predetermined reentry point is reached, the secondary structure is accelerated radially inward using a attitude and orbit control engine at a speed of about 250m/s and into a return trajectory.

The secondary return flow is shown in table 3, the secondary structure is separated from the payload 1, the payload connector 2 is discarded, the secondary main engine 5 is braked to enter a return trajectory, then the inflatable heat shield is unfolded and enters the atmosphere, the recovery parachute 8 at the tail part is unfolded after sufficient deceleration, finally the secondary structure splashes on the sea surface, the buoyancy provided by the inflatable heat shield floats on the sea surface, and the secondary structure is recovered by the recovery ship.

TABLE 2

TABLE 3 Table 3

Preferably, the designed take-off mass of the carrier is 240t, the carrying capacity of the 200kmLEO near-ground track is 12t, and the load coefficient reaches 5%.

Preferably, the air suction mode thrust of the primary engine 13 is 60 tons, the rocket mode thrust is 120t, the mass is 7.5 tons, and the air suction mode thrust-weight ratio is 8. The air suction mode specific impulse is highest 4150s and lowest 2000s, and the rocket mode specific impulse is 430s, as shown in fig. 4.

Preferably, the secondary main engine 5 adopts a 25 ton closed expansion cycle oxyhydrogen engine which is researched, and the engine has the characteristic of high specific impulse, and the specific impulse exceeds 450s. The secondary structure adopts an inflatable heat shield when in use, and has the advantages of light weight, small volume and reusability.

The invention also carries out aerodynamic simulation, firstly, a model can be established by adopting three-dimensional modeling software, then a flow field entity is established, and finally, a grid can be generated and solved by using computational fluid dynamics software to obtain aerodynamic parameters in the flight process of the aircraft.

Fig. 5 a-7 b show trajectory curves, speed and acceleration curves for a mission process.

As shown in fig. 5a and 5b, the maximum overload of the first and second working sections is 2.62g, g is the gravitational acceleration.

As shown in fig. 6a and 6b, the maximum overload during the primary return is 3.19g.

As shown in fig. 7a, 7b, the maximum overload during the secondary return is 7.56g.

As shown by simulation results, the maximum overload of the working section is not more than 3g, and the working section is very stable. The maximum overload of the primary return is slightly above 3g. The secondary return is larger in overload and exceeds 7g due to the ballistic reentry adopted, but is within the bearing range of the existing materials and technologies.

Claims (10)

1. The reusable horizontal take-off and landing two-stage aerospace vehicle is characterized by comprising a primary structure and a secondary structure; the primary structure is an aerospace plane configuration and is used for flushing out the atmosphere and sending the effective load and the secondary structure into the sub-track; the secondary structure is a rocket configuration without lifting surfaces and is symmetrically distributed on the whole and is used for conveying effective loads into a preset track; the primary structure comprises a machine head, a load cabin, an air inlet channel and a primary engine; the load cabin is positioned at the back of the middle section of the machine head and is used for accommodating the secondary structure and the effective load; the air inlet channel is positioned at the front part of the belly and the wing and is used for decelerating air flow and providing air inlet for the primary engine; the primary engine is positioned at the rear part of the belly of the rear section of the aircraft nose and behind the air inlet channel and provides thrust for the power flight section with a primary structure; the secondary structure comprises a secondary main engine, an inflatable heat shield and a recovery parachute, wherein the recovery parachute is used for heat protection when the secondary structure reenters and returns and providing buoyancy, and the recovery parachute is used for final-stage deceleration of the secondary structure.

2. The reusable, level two space carrier of claim 1 wherein the primary structure further comprises wings, movable ducks, vertical tails, front liquid hydrogen tank, rear liquid hydrogen tank, primary liquid oxygen tank; the movable duck wing is positioned on two sides of the aircraft nose and at the front part of the aircraft wing after the middle parts of the two sides of the aircraft nose are deviated, the vertical tail wing is positioned at the tail end of the aircraft wing, the air inlet channel is positioned on the belly and the front part of the aircraft wing, the primary liquid oxygen storage tank is positioned on two sides of the load cabin, the front liquid hydrogen storage tank is positioned in the front cone of the aircraft nose, and the rear liquid hydrogen storage tank is positioned behind the load cabin.

3. The reusable, level lift two space carrier of claim 1, wherein the secondary structure further comprises a secondary liquid oxygen tank, a liquid hydrogen tank, a tank compartment, a payload collar, a attitude and orbit engine, an IVF system; the liquid hydrogen storage tank is arranged at the front, the secondary liquid oxygen storage tank is arranged at the rear, and the liquid hydrogen storage tank is connected with the secondary liquid oxygen storage tank through a tank interval to provide propellant for the secondary main engine; the secondary main engine is arranged in the middle of the secondary liquid oxygen storage tank, the effective load connecting ring is connected with the effective load and the liquid hydrogen storage tank, the attitude and orbit control engine is positioned at the rear part of the liquid oxygen storage tank and used for attitude and orbit control of a secondary structure, the inflatable heat shield is positioned at the inner side of the effective load connecting ring and between the effective load and the liquid hydrogen storage tank, and the recovery parachute is positioned at the rear part of the secondary liquid oxygen storage tank.

4. The reusable horizontal take-off and landing two-stage aerospace vehicle according to claim 1, wherein an adjustable inclined plate is arranged on the inner side of the air inlet channel, and is adjusted according to the flight Mach number, so that the air flow meets the air inlet requirement of the primary engine; when the first-stage engine works in a rocket mode, the air inlet is completely closed by the adjustable inclined plate of the air inlet.

5. The reusable horizontal take-off and landing two-stage aerospace vehicle of claim 1, wherein the primary engine tightly combines three power cycles of turbine, ram, rocket together to achieve full-space, full-speed domain operation.

6. A reusable level lift two space carrier according to claim 1 further comprising nose landing gear, main landing gear, rear landing gear for supporting the level lift two space carrier on a runway.

7. A reusable level lift two space vehicle according to claim 3 wherein the secondary liquid oxygen reservoir is annular.

8. A reusable level lift two space vehicle according to claim 3 wherein the IVF system provides electrical power to the secondary structure, provides gaseous propellant to the attitude and orbit control engine, provides high pressure gas to the inflatable heat shield, and manages the propellant in each tank.

9. A reusable, horizontal take-off and landing, two-stage aerospace vehicle according to claim 1, wherein the primary structure is ignited and then is launched horizontally from the runway; retracting the movable duck wings at 0.9Ma to reduce drag; when reaching 6Ma, switching the first-stage engine from an air suction mode to a rocket mode; when the height reaches 90km, the primary structure is shut down, and the secondary structure and the effective load are discharged from the load cabin; the secondary main engine is then ignited, delivering the payload into the predetermined track.

10. The reusable horizontal take-off and landing secondary aerospace vehicle of claim 1, wherein after separation of primary and secondary structures, the primary structure is reentered into the atmosphere via a short ballistic glide, decelerated by glide, and dropped back to the runway at take-off; the secondary structure is separated from the effective load, the effective load is discarded, the effective load enters a return trajectory through the braking of a secondary main engine, then an inflatable heat shield is unfolded and enters an atmosphere, a recovery parachute at the tail part is unfolded after full deceleration, finally the effective load is splashed on the sea surface, the buoyancy provided by the inflatable heat shield floats on the sea surface, and the effective load is recovered.