RU2841245C1 - Device for providing landing of reusable rocket carrier stage - Google Patents

Abstract

FIELD: astronautics.

SUBSTANCE: invention relates to aerospace engineering, particularly, to devices intended for landing of reusable carrier rocket stage. Device comprises aerodynamic control surfaces, gas-dynamic control system with liquid-propellant power plant (PP) consisting of braking and control engines installed at angle of 15-30°, as well as roll control engines arranged in pairs in transverse plane. Besides, proposed device comprises fuel system consisting of annular fuel and oxidizer tanks with common bearing walls, air-hydraulic system communicated with sustainer tank pressurization system via electrically driven valves and membranes. PP, aerodynamic elements, control equipment and power supply are arranged above any of the tanks of the cruising PP, thus forming an instrument and aggregates compartment.

EFFECT: landing of reusable carrier rocket stage with aerogasdynamic control system.

1 cl, 2 dwg

Description

The invention relates to the field of rocket and space technology, namely in terms of an aerogasdynamic control and landing system for an aircraft, and can be used in the designs of launch vehicle stages, transport spacecraft, and re-entry aircraft flying in the atmosphere over a wide range of speeds.

Currently, a device for an aerodynamic control system for a reusable launch vehicle stage is known based on lattice rudders, which ensures control of the stage during the descent stage in the atmosphere relative to the center of mass along a trajectory close to a ballistic or sliding descent trajectory [1]. The design and layout scheme of the device includes aerodynamic lattice rudders mounted on the body of the aircraft in pairs in mutually perpendicular planes, which ensures 3-channel control of the stage during the descent section, with increased control efficiency due to more complete use of the control surface of the lattice rudder along the flight trajectory. Flight control is performed only during the atmospheric section of the trajectory.

The disadvantage of the device is that in the extra-atmospheric section, control and braking of the stage in the return section is not ensured.

This drawback is partially eliminated in the experimental gliding aircraft HTV-2, developed in the USA by the Defense Advanced Research Projects Agency (DARPA):

- “The Falcon HTV-2 system is experimenting with Mach 20 speeds.” DARPA. Falcon HTV-2 [2];

- Experimental Hypersonic Test Vehicle." 2010 [3];

- «The Falcon HTV-2» [4].

The HTV-2 design uses aerogasdynamic control using deflectable flaps and gas-jet nozzles. The aerogasdynamic control system is based on the use of a split-flap installed on the windward side of the apparatus, which ensures the ability to control the apparatus along the pitch and roll channels. The yaw channel is controlled by a separate gas-jet system with a nozzle block located on the bottom of the apparatus body. Nitrogen was used as the working fluid. The apparatus itself underwent two flight tests in 2010 and 2011 out of three planned as part of the Minotaur IV launch vehicle in an incomplete volume in terms of time and flight range. The aerogasdynamic control system of the apparatus provides only control of the apparatus's motion relative to the center of mass in the atmosphere and does not provide for the braking mode of the apparatus after separation from the upper stage.

An example of a more advanced design of the aerogasdynamic control system is the device used in the experimental transport spacecraft of the European Space Agency, which passed the first flight test on February 11, 2015. Intermediate experimental Vehicle-project IXV (experimental prototype of the spacecraft) [5]. The aerogasdynamic system of this spacecraft is similar to the design implementation in the HTV-2 aircraft: it also uses aerodynamic flaps for control in the atmosphere and a gasdynamic system for flight control after separation from the instrument-assembly compartment in space and partially in the atmosphere for control and stabilization of the yaw channel. When moving in the atmosphere, aerodynamic braking of the spacecraft is carried out, and conditions are formed for the activation of the parachute landing system. The aerogasdynamic control and landing system of the ESA IXV TSC is considered as an analogue.

The closest in technical essence is the control and landing system device for the reusable stage, developed for the launch vehicle in the Falcon 9 FT version [6, 7, 8].

The first stage of the Falcon 9 FT version is specialized in terms of placement of a gas-dynamic orientation and stabilization system during flight in the exoatmospheric portion of the flight and an aerodynamic stage control system using lattice rudders during flight in the atmosphere, as well as installation of a landing device [6, 7].

To ensure landing, the aerogasdynamic control system includes:

1. The reaction control system (RCS), i.e. the actuators of the control, orientation and stabilization system (COS), which is installed in the upper part of the stage body and consists of two blocks with 4 nozzles in each. The control, orientation and stabilization system operates on compressed nitrogen until the lattice rudders open in the atmosphere, and controls the stage on the descent trajectory, including during engine operation in braking mode.

2. Aerodynamic grid rudders (Grid Fins) of the folding type. The rudders are installed on the upper part of the stage in two mutually perpendicular planes in pairs and perform control and stabilization of the stage movement in the atmospheric section of the stage movement after the shutdown of the SUOS engines. In particular, in order to reduce weight, an open hydraulic system is used in the steering drives, which does not require the use of high-pressure pumps.

3. The Merlin ID propulsion system with 9 cruise engines, 8 of which are located around the circumference and one in the center. The special feature of the system is the ability to be launched multiple times on the descent trajectory and use the jet thrust of individual cruise engines to solve the following tasks:

- for braking using the central engine during entry into the atmosphere and soft landing of the stage;

- to slow down the stage upon entering the atmosphere and to possibly slightly reduce the thrust of the propulsion system shortly after the launch vehicle's launch for uniform fuel consumption using three of the nine engines.

Deep throttling of the three Merlin ID thrust engines is achieved by installing a "throttle valve" that regulates the thrust of the engines from 70 to 100% of the nominal thrust, which allows for a flexible thrust control program to be implemented on the trajectory depending on the payload mass and the orbit to which it is to be launched. The thrust of the engines at sea level/in vacuum is 7607/8300 kN, respectively.

4. Navigation equipment for launching the stage to the landing point, which is a control and guidance system to ensure controlled movement from an altitude of H>100 km along a trajectory close to ballistic, with landing at a given point. The control and guidance system is implemented in the SINS version, integrated with the GPS navigation system. According to estimates, its mass can be up to 5 ... 10 kg, taking into account the antenna-feeder device (AFD).

The landing support device with an aerogasdynamic control system for the reusable return stage of the launch vehicle, used on the Falcon 9 launch vehicle version FT, has been adopted as a prototype.

The disadvantage of this device is as follows. On the return trajectory of the stage flight, the control scheme of the center of mass motion is preserved, i.e. the "pushing" scheme, which, unlike the control of the LV motion on the active section, requires more rigid and highly dynamic control to implement controlled motion along the trajectory of a sliding descent in a wide range of speeds: from high supersonic to subsonic under the influence of high-speed and thermal flows in the altitude range from H> 100 to H = 0 km, with landing at a given point with an accuracy of up to 10 ... 15 m (2.7b). This is especially true for the unorganized aerodynamic flow around the stage during a bottom-first descent. In addition, the use of the propulsion system's cruise engines for braking the stage, which uses a multiple-start system with deep throttling of the thrust of individual engines, leads to a loss of specific impulse and an increased consumption of fuel components, since the consumption of fuel components remains consistent with the cruising mode (or close to it) of engine operation. According to data [6, 7, 8], the mass of fuel consumed for braking and control is from 13 to 18% of the total mass of fuel placed in the stage tanks.

The objective of the invention is to develop a device for ensuring the landing of a reusable stage of a launch vehicle with an aerogasdynamic control system, free from the indicated disadvantages.

The required technical result is achieved by the fact that the device is provided with an aerogasdynamic control system, which contains a liquid rocket propulsion system, including two blocks with 4 engines each. Of these, 4 engines are braking and control engines along the pitch and yaw channels. The engines are located in two mutually perpendicular planes so that their longitudinal axes are directed at an angle of 15 ... 30 degrees relative to the longitudinal axis of the stage, and ensure braking of the stage in the pulling mode. The other 4 engines of control along the roll channel are located in pairs diametrically in the transverse plane of the stage. The liquid rocket propulsion system comprises annular-shaped fuel and oxidizer tanks having common load-bearing walls with the load-bearing wall of the stage body and with each other, a pneumatic-hydraulic gas-dynamic control system connected to the pressurization system of the cruise propulsion system tanks and the system for feeding fuel components to the cruise propulsion system tanks by means of solenoid valves and membranes of the fuel component feeding system installed on the pipelines. The liquid propulsion system of gas-dynamic control itself together with the aerodynamic control elements, the control and guidance system equipment, the on-board power source are located between the fuel and oxidizer tanks or above the upper tank of the stage to form the instrument-aggregate compartment of the aerogas-dynamic control and provide extended functional capabilities of such a control scheme.

In such a design variant, the aerogasdynamic system performs a dual function: it ensures both control and stabilization of the stage motion relative to the center of mass along the entire trajectory of the autonomous flight after separation from the launch vehicle, and control of the stage center of mass motion in the linear velocity regulation mode, i.e. in the braking mode. In addition, when controlling the center of mass motion, the stage braking along the entire trajectory is implemented according to the "pull" scheme, bearing in mind that the coordinate of the point of application of the total thrust vector (or the focus of application) X _Fд , created by the gas-dynamic control propulsion system, is located above the coordinate of the stage center of mass X _Tс in a dry or close to empty state, i.e. X _Fд > X _Tст. The positive margin ΔХ _Fд =Х _FДУ - Х _Tст , is (0.3…0.4)L _ст , where L _ст is the stage length, and depends on the installation angle of the gas-dynamic control propulsion system nozzles relative to the longitudinal axis.

Thus, the gas-dynamic control propulsion system enables the return and landing of the stage along the trajectory of a sliding descent. At the same time, the stage's cruise propulsion system performs its only function, namely, the launch of the LV along a given trajectory, and can be unified for both LV variants: with and without stage return.

Structurally, the gas-dynamic control propulsion system is equipped with its own tanks with propellant components, the same as for the main cruise propulsion system, with a fuel consumption calculated to bring the stage to landing along the standard trajectory. In case of deviation from the calculated mode, depending on the mass of the payload and the orbit to which it should be inserted, the pneumatic hydraulic system provides for the supply of propellant components from the stage tanks. For control, stabilization and braking of the stage, the pneumatic hydraulic system of the gas-dynamic control propulsion system provides for the use of multiple-switching solenoid valves with an adjustable switching frequency to gain the required thrust value or total thrust impulse, which are characterized by a minimum aftereffect impulse.

The aerogasdynamic control and landing system, including the aerodynamic control system (drives, control surfaces, etc.), the gasdynamic control propulsion system, fuel tanks, control and guidance system equipment, on-board power sources and landing legs, forms a single structural assembly and is placed in a body installed between the fuel and oxidizer tanks, or in the upper part of the stage, i.e. above the oxidizer tank. In this version, the upper part of the stage with the oxidizer tank and the lower part of the stage with the fuel tank and cruise propulsion system may remain structurally unchanged relative to the expendable stage version, except for technical solutions that ensure increased strength and multiple use as part of the launch vehicle. The same approach remains for the version of placing the aerogasdynamic control compartment in the upper part of the stage, i.e. above the oxidizer tank.

The essence of the invention is explained by a drawing, where Fig. 1 shows a design and layout diagram of the aerogasdynamic control system of the return reusable stage of the launch vehicle, which contains:

1 - stage body;

2 - oxidizer tank;

3 - pipeline;

4 - aerogasdynamic control compartment body;

5 - nozzles of the propulsion system for control and braking;

6 - equipment for the stage control and guidance system with an on-board power source;

7 - annular fuel tank;

8 - toroidal tank of the oxidizer of the gas-dynamic control system;

9 - fuel tank;

10 - power unit compartment;

11 - cruise propulsion system;

12 - antenna insert for satellite navigation equipment;

13 - S/C range insert antenna;

14 - upper oxidizer tank;

15 - lower fuel tank;

16 - power unit compartment;

17 - propulsion system power compensator;

In the upper part of the stage body 1, there is an oxidizer tank 2 with a pipeline 3 for feeding the component to the turbopump unit. In the body 4 of the aerogasdynamic control compartment, there are nozzles of the propulsion system 5 for control and braking, which are installed at an angle of 15°…30° to the longitudinal axis X of the stage, are located in two mutually perpendicular planes of the stage and provide control, orientation, stabilization in the pitch and yaw channels and braking of the stage on the return trajectory or descent trajectory. In the same compartment, there is equipment 6 of the stage control and guidance system with an on-board power source, annular oxidizer tank 7 and toroidal tank 8 of the gasdynamic control system of fuel. In the lower part of the stage, there is a fuel tank 9 and an assembly compartment 10 with a cruise propulsion system 11 of the reusable stage of the launch vehicle. In the upper part of the stage body, there is an antenna insert 12 of the satellite navigation equipment and an antenna insert 13 of the S/C range for transmitting telemetry information (S-range) and for ground and satellite radio communication (C-range). On the body of the aerogasdynamic control compartment, there are control aerodynamic surfaces 14 in the form of rudders or grids. In the transverse plane of the compartment, according to the X-shaped scheme, there are nozzles 15 of the gas-dynamic control system of the roll channel, according to the same scheme, landing legs 16 with carriages and hinged struts of the stage landing system are installed on the compartment body. The center of mass of the stage 17 in a dry or close to empty state is X _t = 0.25 ... 0.30 from the aft cut.

Fig. 2 shows the pneumatic-hydraulic circuit of the gas-dynamic control of the return stage, which includes:

9 - reusable stage fuel tank;

18 - pressure accumulator (PA) for pressurizing the stage tanks and for displacing the components of the stage gas-dynamic control system;

19 - fuel tank pressure accumulator solenoid valve of the stage;

20 - reusable membrane of the fuel supply system of the cruise control unit and the gas-dynamic control unit;

21 - fuel pump solenoid valve of the cruise control unit;

22 - electric valve for supplying fuel to the control unit of the gas-dynamic control and landing system of the reusable stage;

23 - gas generator (GG) of the reusable stage cruise propulsion system;

24 - fuel pump (FP) of the cruise DU of the reusable stage;

25 - oxidizer pump (OP) of the cruise DU of the reusable stage;

26 - reusable disconnecting coupling of the turbopump unit;

27 - turbine (T) of the pumps of the cruise propulsion system components;

28 - pipeline for supplying combustible mixture to the combustion chamber of the cruise engine;

29 - combustion chamber of the cruise engine;

30 - propulsion propulsion nozzle block;

31 - stage oxidizer tank pressure accumulator solenoid valve;

32 - reusable membrane of the oxidizer supply system of the cruise control system and the gas-dynamic control system;

33 - solenoid valve of the cruise oxidizer pump;

34 - solenoid valve for supplying oxidizer to the gas-dynamic control and landing system;

35 - gas-dynamic control system of the roll channel;

36 - solenoid valves for supplying fuel components to the combustion chamber of the gas-dynamic control system of the roll channel;

37 - electric valves for feeding fuel components into the combustion chamber of the gas-dynamic control system, orientation, stabilization by pitch and yaw channels and braking of the stage.

The design-layout diagram and the pneumatic-hydraulic diagram of the aerogasdynamic control system of the reusable stage of the launch vehicle shown in Figs. 1 and 2 were obtained based on the results of design studies with assessments of the design features of the reusable stage in various variants. The assessments were carried out for variants implementing:

- minimum energy costs, meaning fuel costs for the implementation of the return trajectory. The gas-dynamic system performs a dual function: control and stabilization of the stage motion relative to the center of mass along the entire trajectory of the autonomous flight from the moment of separation from the launch vehicle and control of the motion of the center of mass of the stage in the linear velocity control mode, i.e. in the braking mode. The braking process is performed after orientation and stabilization of the stage position on the trajectory at altitudes from the maximum achievable H ≈ 120 km to H = 40 ... 50 km, at which the maximum specific impulse of the liquid-propellant rocket engine is realized and the stage flight speed is reduced from V _вx ≥ 1300 m / s to transonic and a transition to aerodynamic control of the stage in dense layers of the atmosphere is possible. Throughout the trajectory, the control of the stage motion relative to the center of mass and the motion of the center of mass of the stage is implemented according to the normal and "pull" scheme, bearing in mind that the point of application of the total thrust vector (or the focus of application) created by the gas-dynamic control PS is located above the center of mass of the stage in a dry or close to empty state, i.e. X _FDU - X _Tst > 0, and ΔX _FDU is (0.3…0.4)L _CT relative to the aft section of the stage and depending on the installation angle of the gas-dynamic PS nozzles relative to the longitudinal axis. In the declared configuration, the focus of application of the total thrust vector is determined by the installation angle of the gas-dynamic PS nozzles relative to the longitudinal axis ϕ, which can be structurally selected in the range of 5°…30°;

- a combined gas-dynamic control propulsion system implementing both stage braking on the return trajectory in landing mode and stage control, orientation and stabilization relative to the center of mass. The pneumohydraulic circuit of the gas-dynamic control propulsion system is based on a displacement system for feeding fuel components to the combustion chamber and has its own fuel component tanks, the same as those on the cruise propulsion system. The GU propulsion system itself consists of 4 high-thrust engines providing stage braking and landing and stage control and stabilization via pitch and yaw channels, and 4 low-thrust engines for stage control and stabilization via the roll channel. According to estimates, the thrust of each high-thrust engine is ~ 200…250 kg, and the thrust of a low-thrust engine is ~ 40…50 kg for ultra-light or light class launch vehicles. The GU DU is equipped with fast-acting solenoid valves, for example, of the pulse type, for feeding fuel components into the combustion chamber with a minimum after-effect pulse. The displacement system for feeding fuel components into the combustion chamber of the gas-dynamic control DU is integrated with the fuel supply system of the cruise DU with the possibility of their separate stage-by-stage operation. Upon completion of the operation of the cruise stage DU, the pneumohydraulic circuit provides for its shutdown, and only a part of the pneumohydraulic circuit of the GU DU functions. When the calculated volume of fuel components in the GU DU tanks is consumed, the pneumohydraulic circuit provides for their replenishment from the stage tanks;

- minimal design modification of the stage when switching to a reusable mode of use. This becomes possible due to the fact that the cruise propulsion system does not participate in the stage landing process. This creates the possibility of unifying the stages of single-use and reusable use on LVs of any mass class. The only difference is that the design of the first stage must be adapted for reusability in both cases. This includes, first of all, the reusable cruise propulsion system, fastening devices for those systems and devices that must be on the reusable stage option. Among them, for example, fastening the landing device on the fuel tank, devices for transporting the stage to the place of scheduled maintenance, as well as strengthening its power structure;

- the possibility of autonomous flight testing of the aerogasdynamic control system, as a structural element of the stage, on the return trajectory of descent.

The presented results confirm the increased efficiency of using the developed aerogasdynamic control system for the reusable return stage based on the proposed design-layout and pneumohydraulic schemes with an aggregate intertank compartment of the aerogasdynamic control. The use of such a system creates opportunities for implementing options for constructing a reusable launch vehicle stage with improved flight performance characteristics in terms of implementing a flight along a sliding descent trajectory, i.e. along a trajectory with low aerodynamic quality K = 0.1 ... 0.15, with landing at a given point with minimal energy costs while maintaining its mass characteristics.

The device operates as follows. At the moment of separation of stage 1 from the LV, the power supply system of the cruise propulsion system is blocked by closing the electric valves 21 and 33 with disconnection of the TNA 27 of the GG 23 power supply and the inertial control system of stage 6 is switched on. Upon its command, the aerogasdynamic control system is activated by opening the electric valves 22 and 34 of the fuel and oxidizer supply of the propulsion system of the gasdynamic control and the drives of the aerodynamic control bodies 14. In this case, the pressurization system AD 18 of the fuel tanks 9 and oxidizer 2 with valves 19 and 31, as well as the reusable membranes 20 and 32, respectively, remain open.

During the inertial flight using the propulsion system consisting of engines 5 and 35, the stage is oriented and stabilized relative to the center of mass. Upon reaching the maximum altitude H = 100 ... 150 km, the stage is turned onto the landing trajectory and then braked using 4 engines 5 in the controlled braking mode. Braking is performed with the calculation of reducing the flight speed to 1200 ... 1400 m / s at an altitude of 40 ... 45 km to ensure the movement of the stage in the dense layers of the atmosphere at an acceptable speed. At this stage, the aerodynamic controls and stabilization of the stage's movement along the pitch, yaw and roll channels open and begin their work with the simultaneous operation of engines 5 in the braking mode. Depending on the orbit parameters and the mass of the LV payload forming the return flight trajectory and, as a consequence, the possible increased fuel consumption for braking and landing, the fuel from the tanks of the cruise stage, when the solenoid valves 21 and 33 are closed, can be pumped into the tanks of the gas-dynamic control and landing propulsion system, and then to the braking, control and stabilization engines 5 and 37 via the pitch and yaw channels and to the control and stabilization engines 4 and 35 via the roll channel. By the time of landing in the terminal control mode, the speed of the stage is brought to the permissible value of 2 ... 4 m / s. The landing gear struts 16 open before the stage touches the surface of the landing pad.

Thus, the new technical result is achieved by the fact that a technology is proposed for ensuring the landing of a stage based on an aerogasdynamic control system, in which the gasdynamic control system, aerodynamic controls, control and guidance system equipment and on-board power sources are arranged in a single assembly, forming an instrument-aggregate compartment of the aerogasdynamic control. A pulling control scheme for the movement of the center of mass on the stage descent trajectory and a normal control scheme for movement relative to the center of mass of the stage are implemented, while maintaining the overall mass characteristics of the stage relative to the prototype. In the design and layout scheme of the stage, the aerogasdynamic control compartment can be located both in the intertank or above-tank volume of the stage, and separately above the stage, depending on the adopted pneumatic hydraulic system, forming an autonomous landing unit of the returning stage.

Based on the proposed technology, it becomes possible to conduct autonomous flight testing of the aerogasdynamic control system as part of the demonstrator of the return stage, which ensures an increase in the level of technology readiness in the process of developing the return stage in a shorter time frame.

SOURCES OF INFORMATION

1. Patent for invention No. 2800531 dated July 24, 2023. “Device of an aerodynamic control system for a reusable return stage of a launch vehicle.”

2. https://www.google.com/imgres?imgurl=https://www.globalsecurity.org/space/systems/images/htv-3-imagel.jpg&imgrefurl.

3. https://www.militaryfactory.com/aircraft/detail.php?aircraft_id=885.

4. https://images,app.goo.gl/LZyNTHaCjbJ4gUxK9.

5. http://en.Wikipedia.org/wiki/Intermediate experimentalVehicle.

6. https://ru.wikipedia. org/wiki/Falcon_9.

7. https://en.wikipedia. org/wiki/Falcon_9 Full Thrust.

8. US 8,678,321. - B2 (45). - Date of Patent: Mar. 25, 2014.

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10. Akhmetshin K.Sh., Kiryukhin S.Yu., Ryabinin A.S. Comparative analysis of methods for regulating the thrust of liquid and solid rocket engines. Siberian State Agrarian University named after academician. M.F. Reshetnev. Reshetnev readings. 2013. Pp. 110-111. Electron, Wikipedia resource. Electron, data. - Access mode: https://cyberleninka.ru/article/n/sravnitelnvv-analiz-sposobov-regulirovaniya-tvagi-zhidkostnyh-i-tverdotoplivnvh-raketnyh-dvigateley/viewer

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Claims (1)

A device for ensuring the landing of a reusable stage of a launch vehicle, comprising aerodynamic control elements in the form of control surfaces, a gas-dynamic control system relative to the center of mass of the stage, characterized in that a liquid propulsion system is introduced into the gas-dynamic control system, consisting of 4 braking and control engines in the pitch and yaw channels located in two mutually perpendicular planes and the longitudinal axes of which are directed at an angle of 15 ... 30 degrees relative to the longitudinal axis of the stage with ensuring braking of the stage in the pulling mode, and 4 control engines in the roll channel, paired diametrically located in the transverse plane of the stage, containing annular-shaped fuel and oxidizer tanks having common load-bearing walls with the load-bearing wall of the stage body and with each other, a pneumatic hydraulic system connected to the pressurization system of the tanks of the cruise propulsion system and a system for feeding fuel components to the tanks of the gas-dynamic control propulsion system by means of installed on the pipelines of the solenoid valves and membranes of the fuel component supply system and thrust vector control, and the propulsion system itself, together with the aerodynamic control elements, the control and guidance system equipment and the on-board power source, is located above the tank of one of the fuel components of the stage's cruise propulsion system, forming the instrument and unit compartment of the aerogasdynamic control.