50 results about "Air breathing" patented technology

The invention discloses a guide wall for airplane ground test run, which comprises a U-shaped barrier part enclosing an airplane, the barrier part comprises a side guide wall corresponding to the side of the airplane and a rear guide wall corresponding to the tail of the airplane, the side guide wall comprises a windward side and a leeward side, and the leeward side is connected with a fixing frame. The fixing frame is connected with a ground anchor bolt to support the side flow guide wall, and the rear flow guide wall is fixedly connected with ground concrete; the length of the side flow guide wall exceeds the position of the nose during the test run of the airplane, and an air inlet channel is arranged at the side flow guide wall corresponding to the side wing during the test run of the airplane, penetrates through the side flow guide wall and supplements airflow to the engine close to the side wing of the air inlet channel; by arranging the air inlet channel, air flow is introduced from the side face of the aircraft directly facing the engines, the problem that the test is affected by insufficient air suction of the double engines due to the fact that the side face flow guide wall is far longer than the fuselage is solved, the internal and external pressure difference of the flow guide wall is reduced, and test run of the aircraft is facilitated.

A vessel for transporting cargo by water comprises one or more propulsion devices with at least one air-breathing engine from the one or more propulsion devices disposed entirely above a surface of the water in a high-speed mode of operation. The least one air-breathing engine is configured to bleed airflow into a fluid circuit. The fluid circuit routes the bleed airflow to one or more fluid outlets disposed entirely below the surface of the water in the high-speed mode of operation. At least some of the one or more fluid outlets release the bleed airflow in the high-speed mode of operation generate a boundary layer along at least a portion of a hull of the vessel to reduce a skin friction drag thereof.

The invention relates to an air suction type cusped field electric propulsion system with multiple working medium supply modes and a supply method. The invention relates to the technical field of air suction type electric propulsion systems. The system comprises an electric thruster and collection device module, a microwave cathode module, a xenon storage and supply module and a control and energy supply module. According to the air-breathing cusped field electric propulsion system with the multiple working medium supply modes, the adaptive capacity of the electric thruster under the complex working condition is improved through the xenon working medium carried by the air-breathing cusped field electric propulsion system while the ultra-low orbit rarefied atmospheric environment is fully utilized. In addition to an ultra-low rail air-breathing type electric propulsion task, the xenon working medium carried by the ultra-low rail air-breathing type electric propulsion device can enable the ultra-low rail air-breathing type electric propulsion device to give consideration to a higher rail range, and a traditional xenon electric propulsion rail interval and an ultra-low rail space of 150 km or above are covered.

A self-contained vertical takeoff and landing vehicle (VTOL) with aspect ratio of height to width no greater than 5.3:1 and with both rocket and airbreathing propulsion systems that can selectively operate one at a time or both at the same time and fuel tanks that have internal lattice structures formed of groups of interconnected multiplicities of fuel cells such as gyroids that line and are integrated with a vehicle skin to form a structural skeleton strengthened with said lattice structures that also serve to reduce fuel slosh compared with fuel tanks lacking such lattice structures and provide other benefits such as fuel temperature profile control.

This invention provides an air-breathing electric propulsion air chamber, including a grid, an air inlet, and a rectifying device. The bottom of the grid is connected to the front end of the air inlet, the inner wall of the air inlet has a parabolic structure, and the rear end of the air inlet is connected to a long tube. The rectifying device is mounted inside the air inlet via a support frame. The rectifying device has a convex mirror-like structure, with a parabolic inner wall. Its focal point is the same as the focal point of the parabolic structure on the inner wall of the air inlet, and the opening of the rectifying device faces in the opposite direction to the opening of the air inlet. This invention refracts the incoming gas twice, finally injecting it parallel into the long tube at the rear end of the air chamber. This reduces the collision and corrosion of the long tube wall by air particles, thereby greatly improving the lifespan and performance of the air-breathing electric propulsion system and increasing the gas collection rate, providing a new technical approach for the development of air-breathing electric propulsion.

The application provides a kind of collaborative jet airfoil optimization design method based on combined parameterization, belongs to the field of aircraft aerodynamic design and flow control.This method firstly carries out CST parameterization to the original basic airfoil for constructing original collaborative jet airfoil;With blowing / air suction port position, size as design variable, the curve of subsidence section is fitted by original basic airfoil linear interpolation, realizes the adaptive association of subsidence section and original basic airfoil.Subsequently, a joint design space containing CST parameterization variable, collaborative jet airfoil design variable and jet control parameter is constructed, the initial proxy model is constructed and updated to convergence by using SBMO framework.The application carries out global collaborative optimization to the subsidence section channel geometry and the overall profile of original basic airfoil, solves the parameter coupling problem of the basic geometry shape of collaborative jet airfoil and subsidence section, and can significantly improve the aerodynamic performance of collaborative jet airfoil.

The application discloses a combined engine and a control method thereof. The combined engine comprises a control assembly, an air inlet channel, a turbine core engine, a nuclear reactor system, a propellant supply system and a nozzle assembly which are connected through pipelines. The control assembly is used for controlling the air inlet end of the nuclear reactor system to be selectively connected with the air inlet channel, the turbine core engine or the propellant supply system through the pipelines. The air outlet end of the nuclear reactor system is connected with the nozzle assembly. The air inlet end of the nuclear reactor system is connected with the air inlet channel, the turbine core engine and the propellant supply system through the pipelines respectively by the control assembly, so that the combined engine enters the nuclear ram mode, the nuclear turbine mode and the non-suction mode respectively. The combined engine has the characteristics of high fuel economy, working in the atmosphere-free environment, realizing different speed domain flights by using different engine modes and meeting the needs of space transportation of the super-speed aircraft.

The application discloses an ultra-low orbit adsorption type air-breathing pressurization system, which is used for an ultra-low orbit electric propulsion system and comprises an adsorption pressurizer on one side and a gas storage on the other side; the adsorption pressurizer is used for adsorbing rare thin gas molecules under low-temperature conditions and releasing the gas molecules under high-temperature conditions, and generates a pressurization effect; and the gas storage is used for storing the pressurized gas molecules and releasing the gas molecules on demand. The application discards the traditional mechanical compression mode and innovatively adopts an adsorption type pressurization principle. Through integrated design of a rotating staggered valve and an adsorption plate group, gas guiding, flow channel switching and adsorption pressurization functions are integrated into one, forming a highly compact overall structure. The application discards the traditional mechanical compression mode and innovatively adopts the adsorption type pressurization principle. The problems of a complex structure, a large size, extremely high power consumption and prominent energy contradiction of the prior art ultra-low orbit electric propulsion system are solved.

This invention belongs to the field of plasma flow control technology for aircraft, specifically relating to an air-breathing plasma jet generator suitable for active flow control of aircraft. It is installed within an internal cavity of the aircraft, connected to an air intake and a nozzle. The air-breathing plasma jet generator generates orthogonal electric and magnetic fields. The electric field includes components in the direction of airflow and perpendicular to airflow. The magnetic field is used to confine electrons to gyroscopic motion and to ionize them through collisions with gas molecules. Furthermore, the magnetic field causes the ion current perpendicular to the airflow direction to generate an Ampere force in the direction of airflow under the influence of the magnetic field. This invention effectively solves the problem of exciter failure during high-speed flight by embedding the exciter entirely inside the aircraft and combining an innovative design of integrated excitation using magnetic fields, thermal energy, and acoustic waves, significantly improving the speed applicability range of plasma flow control technology.

The invention relates to the technical field of air intake systems of aero-engines, and discloses a self-adaptive hybrid air suction system of an aero-piston engine, which comprises an engine body for providing power for aviation flight, and is characterized in that an air suction unit for air intake is fixedly arranged at one end of an inner cavity of the engine body; the air suction unit comprises an air suction part used for air suction, a vortex generator used for improving air mixing and an air inlet adjusting mechanism used for adjusting the air inlet angle, the air inlet adjusting mechanism is fixedly connected with the air suction part through the vortex generator, and an air compression unit used for compressing air is fixedly arranged at one end of the air suction unit; the end, away from the air suction unit, of the air compression unit is fixedly connected with a combustion unit used for combusting compressed air, the end, away from the air compression unit, of the combustion unit is fixedly connected with an exhaust unit used for exhausting air to provide power, the air inlet mechanism can be adjusted to dynamically adapt to the flight working condition, the air inlet efficiency and stability are improved, and the response speed is increased; and the combustion efficiency is optimized.

The application discloses a kind of inner and outer flow integrated high-speed full-flow surface waverider aircraft design method, it is related to aerodynamic shape design field, comprising: S1, based on the constraint condition of incoming flow condition and axial symmetry curved surface cone, utilize characteristic line method to obtain two mutually matched first axial symmetry curved surface cone flow field, second axial symmetry curved surface cone flow field;S2, in two axial symmetry curved surface cone flow field, respectively construct upper waverider surface, lower waverider surface, precursor inlet of waverider aerodynamic configuration using streamer tracking method;S3, upper waverider surface, lower waverider surface and precursor inlet are assembled according to flow field characteristics, and the appearance structure of inner and outer flow integrated high-speed full-flow surface waverider aircraft is obtained.The application realizes the design of high-speed air-breathing waverider configuration from the level of theoretical method, all components are designed using flow surface and the flow field between each component is matched, which provides a very effective and efficient method and technical support for designing waverider configuration meeting engineering practical requirements.

The integrated ultra-low orbit constellation carrier platform comprises an air inlet cabin section, a middle cabin section and a tail cabin section, the air inlet cabin section comprises an air inlet, the middle cabin section comprises a radio frequency ionization chamber, an electric control solid propeller storage cabin and a satellite storage cabin, the tail cabin section comprises a magnetic jet acceleration pipe, a propellant supplement mechanism and an electric control solid rocket upper stage, the air inlet, the radio frequency ionization chamber and the magnetic jet acceleration pipe are connected to form an air-breathing electric propulsion system of the carrier platform. The satellite is connected with the corresponding electric control solid rocket upper stage in a launch channel. In the satellite transportation into orbit networking stage, the electric control solid rocket upper stage carrying the satellite is ejected from the cabin, the electric control solid rocket upper stage uses electric control solid propellant as a propellant working medium and can be repeatedly ignited and started, after reaching a target orbit, the satellite is separated from the electric control solid rocket upper stage and is unfolded and networked by itself, and the electric control solid rocket upper stage returns to the carrier platform for fuel supplement by relying on the remaining fuel, and is ready to start the next satellite transportation task.

The invention relates to the technical field of spacecraft electric propulsion, in particular to an air suction type electric propulsion system turbo molecular pump supercharging performance experimental verification system which comprises a vacuum cabin, a molecular pump, a closed cavity, a gas cylinder, a discharge chamber simulation cavity, a backing pump and a vacuum gauge. The gas cylinder is communicated with the closed cavity through a first gas conveying pipeline; the output end of the molecular pump is communicated with the gas inlet of the discharge chamber simulation cavity through a second gas transmission pipeline; a gas outlet of the discharge chamber simulation cavity is communicated with the interior of the vacuum chamber through a third gas transmission pipeline; and the backing pump is communicated with the second gas conveying pipeline through a gas pipeline. According to the application, the pressure environment condition required at the inlet of the turbine can be accurately created, the working condition of the molecular pump when the air-breathing electric propulsion system works can be accurately simulated, the supercharging performance of the molecular pump is subjected to experimental verification, and the feasibility of the molecular pump for air-breathing active supercharging can be evaluated.

Disclosed is an aircraft having a rocket-based combined cycle (RBCC) propulsion system including a rotating detonation rocket engine (RDRE) and an airbreathing jet engine. The RDRE is located within a throat area of the jet engine and is configured to be cooled by air bypassed from upstream of the jet engine.

This invention relates to a multi-propellant supply mode air-breathing tangential field electric propulsion system and its supply method. The invention pertains to the technical field of air-breathing electric propulsion systems. The system includes: an electric thruster and collection device module, a microwave cathode module, a xenon gas storage and supply module, and a control and energy supply module. The multi-propellant supply mode air-breathing tangential field electric propulsion system provided by this invention, in addition to fully utilizing the thin atmosphere of the ultra-low Earth orbit (ULE), enhances the adaptability of the electric thruster under complex operating conditions by utilizing its own xenon propellant. Beyond ULE air-breathing electric propulsion missions, the xenon propellant it carries enables it to operate over a wider orbital range, covering the orbital range of traditional xenon electric propulsion and ULE space above 150 km.

The invention relates to an air-breathing electric propulsion system ground test system and method, and belongs to the field of ultra-low orbit space electric propulsion, and the system comprises a spiral wave plasma source, an incoming flow parameter measurement device, a tested air-breathing electric propulsion system, a micro-thrust frame, a vacuum system, a gas cylinder and a micro-thrust frame measurement system. A gas cylinder is used as a gas source to generate simulated incoming flow through a spiral wave plasma source, the simulated incoming flow enters a tested air-breathing electric propulsion system, the thrust of the tested air-breathing electric propulsion system is calculated by measuring net thrust, and then specific impulse is calculated. The spiral wave plasma source is used for generating rarefied gas plasma incoming flow equivalent to neutral gas incoming flow, incoming flow with different track heights and different component contents can be simulated by adjusting flow, power, magnetic field intensity and the like, and performance verification of the air suction type electric propulsion system is achieved by measuring net thrust.

The invention belongs to the technical field of aircrafts, and aims to solve the problems of resistance divergence effect and difficult engine air intake of a traditional aircraft. The air-breathing high-subsonic-velocity aircraft comprises a nose, a fuselage, a tail, wings and an empennage, the nose is connected with the tail through the fuselage, the wings and the empennage are connected with the fuselage, the wings are supercritical airfoil type sweepback wings, an embedded air inlet channel is formed in the lower surface of the tail and communicated with an engine in the tail, and the engine is connected with the tail through the embedded air inlet channel. The embedded air inlet channel can introduce air on the lower surface of the tail into the engine so as to meet the air inlet requirement of the engine. According to the aircraft, the wings are arranged to be the supercritical wing type sweepback wings, through the unique geometric shape and sweepback arrangement, the resistance divergence phenomenon is effectively delayed, the effective speed of airflow perpendicular to the front edge is reduced, accordingly, shock wave generation is delayed, an embedded type air inlet channel is formed in the lower surface of the tail, and the impact resistance of the aircraft is improved. The air inlet requirement of an engine of an aircraft is met, and meanwhile the flight resistance and the radar cross section are reduced.

The invention belongs to the technical field of design of air inlet systems of aviation aircrafts, and particularly relates to a subsonic air inlet channel self-adaptive boundary layer suction and removal device based on differential pressure driving. The device is characterized in that a micropore suction component is arranged on the surface of a fuselage skin in front of an air inlet lip to form an air suction micropore area, and the air suction micropore area is an airflow stagnation high-static-pressure area and is used for sucking airflow of a boundary layer; the gas collection pressure stabilizing cavity is arranged on the inner side of the micropore suction component in a sealing manner and covers all the suction micropore areas; the exhaust assembly is arranged in a low-static-pressure area on the surface of the fuselage or the surface of the wing, and the surface static pressure coefficient of the low-static-pressure area in a subsonic aircraft state is lower than that of an area in front of an air inlet lip; one end of the flow guide channel is communicated with the gas collection pressure stabilizing cavity, and the other end is communicated with the exhaust component to form a communicated airflow channel. The device has the advantages of being simple in structure, light in weight, capable of adjusting the air suction amount in a self-adaptive mode and the like, and the stealth performance and the total pressure recovery coefficient of the air inlet channel are remarkably improved.

The application belongs to the technical field of pulse combustion wind tunnel test, and discloses a hydrogen-oxygen combustion heater applied to a high-temperature air inflow pulse combustion wind tunnel. The hydrogen-oxygen combustion heater of the pulse combustion wind tunnel comprises, from front to back, a high-temperature flow-limiting throat, a mixing section, an injector, a combustion chamber and a nozzle; the high-temperature flow-limiting throat comprises, in sequence, a contraction section, a throat section and an expansion section, the expansion section is provided with an oxygen gas collecting cavity and a plurality of oxygen gas injection holes; the injector is composed of an injection unit, a front cover plate and a rear cover plate; a plurality of injection units distributed in a central symmetry mode penetrate the front cover plate and the rear cover plate from front to back; the cavity surrounded by the front cover plate, the rear cover plate and the injection unit is a hydrogen gas collecting cavity; the inner cavity channel of the injection unit is first contracted and then expanded, and the expansion section is provided with a plurality of hydrogen gas injection holes distributed uniformly in a circumferential direction. The effective mixing of high-temperature air, hydrogen and oxygen is realized, self-ignition is achieved, and high-temperature and high-pressure gas generated thereby is used to carry out wind tunnel test of air-breathing hypersonic aircraft.

The invention relates to a method and a device for evaluating the starting capability of a high-speed air inlet duct by utilizing protective cover separation. The method comprises the following steps: constructing a background grid for an overall computational domain of an air-breathing aircraft; overlapping grid blocks are constructed for the protection cover; peripheral boundaries of the overlapped grid blocks are defined as overlapped boundaries; the overlapped grid blocks and the background grids are assembled; performing flow field initialization on the overall computational domain according to the parameters of the target working condition, and simulating a supersonic flow field; simulating rotary cover opening motion of the protective cover and solving an unsteady flow field to obtain an initial flow field; when the protective cover rotates to a preset cover opening angle, the protective cover continues to move in the direction away from the air-breathing aircraft until the protective cover moves to a far-field position; and deleting the overlapped grid blocks from the overall computational domain, continuously monitoring the air inlet flow state evolution of the high-speed air inlet in the overall computational domain by adopting an unsteady numerical method based on the obtained initial flow field, and evaluating the starting capability of the high-speed air inlet.

The invention belongs to the technical field of engine parameter calculation, and particularly relates to an air suction cruise hypersonic aircraft nozzle parameter equivalent estimation method, which comprises the following steps of: enabling an air inlet channel, an isolation section and a combustion chamber of a scramjet engine to be equivalent to a combustion chamber of a rocket engine; the contraction ratio of the jet pipe of the scramjet engine is equivalent to the expansion ratio of the jet pipe of the rocket engine; the method comprises the following steps: enabling air flow and fuel injection flow entering a combustion chamber of the scramjet engine to be equivalent to premixed gas which is uniformly mixed according to a preset proportion, and determining formula parameters of an equivalent propellant based on the premixed gas; converting the flow field parameters at the air inlet of the scramjet engine into the total temperature and the total pressure of a combustion chamber of an equivalent rocket engine; calculating nozzle parameters of the scramjet engine by adopting a thermodynamic calculation method based on the parameters obtained in the above steps; the problems that an existing scramjet engine nozzle parameter is large in solving difficulty and high in calculation cost are solved.

This invention relates to a method and apparatus for evaluating the start-up capability of a high-speed air intake using a protective shield separation. The method includes: constructing a background mesh for the overall computational domain of an air-breathing aircraft; constructing overlapping mesh blocks for the protective shield; defining the four boundaries of the overlapping mesh blocks as overlapping boundaries; assembling the overlapping mesh blocks with the background mesh; initializing the flow field of the overall computational domain according to the parameters of the target operating condition, simulating a supersonic flow field; simulating the rotational opening motion of the protective shield and solving for an unsteady flow field to obtain an initial flow field; after the protective shield rotates to a predetermined opening angle, allowing it to continue moving away from the air-breathing aircraft until the protective shield reaches a far-field position; deleting the overlapping mesh blocks from the overall computational domain, and continuously monitoring the evolution of the air intake flow state in the overall computational domain using an unsteady numerical method based on the obtained initial flow field, thereby evaluating the start-up capability of the high-speed air intake.

The application discloses an air turbine rocket engine, which comprises a pulse detonation system and a air-breathing engine system with a compressor and a mixing chamber, the pulse detonation system is provided with a pulse detonation combustion chamber, a gas turbine and a rich fuel gas delivery pipeline, the gas turbine is connected with the compressor, and the rich fuel gas delivery pipeline is communicated with the pulse detonation combustion chamber and the mixing chamber. The air turbine rocket engine utilizes the characteristics of high temperature and high pressure of the outlet gas of the pulse detonation combustion chamber to realize high-efficiency work of the gas turbine.

The application provides a novel gas collecting device and an air-breathing electric propulsion system, comprising a rotary wing and an air-breathing nozzle, the air-breathing nozzle is provided with an air inlet and an air outlet at two ends respectively, the radius of the air inlet is larger than that of the air outlet, the rotary wing is arranged on the inner wall of the air-breathing nozzle close to the air inlet; one end of the rotary wing is connected with the air-breathing nozzle, the other end of the rotary wing is arranged obliquely towards the air outlet, and there is an included angle β between the length direction of the rotary wing and the air-breathing nozzle. The application has simple structure, the number of particles in positive motion can be greatly increased through the change of the gas flow in the air-breathing nozzle by the rotary wing, the gas collecting rate of the device is improved, and the rotary wing is not contradictory to the existing strategies of "adding a rhombic cone" and "adding a front-end grid", so that multiple optimization strategies containing the rotary wing can be used at the same time, and the collecting rate of the air-breathing device is improved.

The application discloses an atomic oxygen conversion method and an air-breathing electric propulsion filtering system. The atomic oxygen is ionized and then recombined to generate oxygen, so that high-speed and controllable oxygen conversion is realized, and the reaction efficiency is high. The air-breathing electric propulsion filtering system comprises an air inlet section, a contraction section and a straight section. The air-breathing electric propulsion filtering system is provided with a central electrode and a side wall electrode, so that the atomic oxygen in the super-low orbit atmosphere is pre-ionized, and is pushed to a side wall thickening area under the action of an electric field. The atomic oxygen ion obtains electrons from the metal material of the side wall thickening area, and then re-generates atomic oxygen. A high-concentration atomic oxygen aggregation area is generated, so that the recombination rate of the atomic oxygen is greatly improved. The generated oxygen molecules are rectified by a rectifying mechanism, and then replace the original strong corrosive atomic oxygen to enter an ion thruster, so that the corrosion of the atomic oxygen in the super-low orbit atmosphere to key components of the ion thruster is effectively avoided.

The invention discloses an air-breathing type electromagnetic propulsion device for an ultra-low orbit satellite, and relates to the technical field of electromagnetic acceleration propulsion. In the device, an air-breathing type gas collecting assembly collects and compresses mixed gas of an ultra-low orbit; the radio frequency plasma ionization assembly ionizes the composite working medium gas to form plasma; the plasma heating acceleration assembly injects energy of a high-power radio frequency power supply into plasma through electromagnetic wave coupling, and heating and acceleration of the plasma are achieved. The vector thrust control assembly restrains and controls the heated and accelerated plasma to form a directional ion beam; the electric propulsion main control assembly collects key state parameters of the air suction type gas collection assembly, the radio frequency plasma ionization assembly, the plasma heating acceleration assembly and the vector thrust control assembly through sensors, and closed-loop control is conducted on the four assemblies according to received or preset control instructions and in combination with the key state parameters. Efficient ionization can be achieved under low working medium pressure, and wide-load efficient and stable operation can be achieved.

The invention provides a flexible composite film structure for a pressure stabilizing cavity of an air-breathing electric propulsion system and a preparation method, and relates to the technical field of spacecraft propulsion system materials and structures. Aiming at the requirements of low-orbit spacecraft components on stable air pressure control, light weight and space environment protection, an integrated solution that a pressure stabilizing cavity main body adopts a four-layer composite film structure of a flexible substrate, a barrier layer, a functional layer and an atomic oxygen prevention layer is provided. The flexible substrate selects polyimide to provide basic bearing and flexibility, and the barrier layer adopts a direct-current magnetron sputtering technology to deposit an aluminum film on the substrate, so that gas trapping sealing and leakage barrier are realized. The functional layer is made of carbon fiber or carbon nanotube reinforced polyimide and is formed through solution spraying and curing, the overall mechanical strength of the composite film is improved, and heat dissipation is assisted. The atomic oxygen prevention layer adopts a plasma-assisted enhanced chemical vapor deposition technology to realize preparation of a siloxane film and is used for resisting erosion of atomic oxygen in a space environment and prolonging the service life of the composite film.

The application discloses an electric propulsion system, an electric propulsion gas supply device and a gas supply method. The electric propulsion gas supply device comprises a control module, a low-temperature gas suction module, a refrigeration module and a gas storage module. The low-temperature gas suction module is used for sublimating gas into solid working medium after contacting with the gas in the space environment, and also converts the solid working medium into fluid working medium into the gas storage module. The refrigeration module is used for refrigerating the low-temperature gas suction module to make the low-temperature gas suction module sublimate the contacted gas into solid working medium. The control module is used for controlling the refrigeration module to stop refrigerating the low-temperature gas suction module when the solid working medium in the low-temperature gas suction module is greater than a preset value, and controlling the low-temperature gas suction module to convert the solid working medium in the low-temperature gas suction module into fluid working medium. The gas storage module is used for storing the fluid working medium. The application can absorb the gas in the low orbit, can not carry working medium or carry a small amount of working medium, can prolong the on-orbit operation time of the satellite, and can also ensure the accurate control of the operation attitude and operation track of the satellite.

The present application relates to a kind of radial connecting device of air-breathing hypersonic missile and engine, belong to the field of missile structure.The present application includes: elastic support device (2), adjusting screw (3), support mounting seat (5), locking mechanism;The locking mechanism includes: locking nut (4), locking screw (6);The elastic support device (2) includes: support leaf spring (7), heat sink leaf spring (8);The adjusting screw (3) is vertically inserted into support mounting seat (5), locking nut (4) is provided at the contact place of adjusting screw (3) outside support mounting seat (5), the locking nut (4) is around and is provided in adjusting screw (3) outside, can adapt to the circumferential and radial thermal deformation caused by engine when rocket cruise under the requirement of meeting support rigidity.

The invention belongs to the field of wing trailing edge flow control, and particularly relates to an airfoil fusion air suction type wing trailing edge flow control device. The device comprises a wing, the wing is provided with a wing surface hollow interlayer, and an airflow flowing channel from the front edge of the wing to the rear edge of the wing is formed in the wing surface hollow interlayer; the wing trailing edge control surface is arranged on the wing trailing edge; the grating type air entraining opening is formed in the upper surface of the front edge of the wing; the flow controller is arranged in the air flow flowing channel; the grating type air outlet is formed in the upper surface of the rear edge of the wing, and airflow flowing out of the grating type air outlet acts on the control surface of the rear edge of the wing. The external flow field can be improved and adjusted, and the effects of improving the aerodynamic characteristics of the aircraft, improving the control efficiency and enhancing the flight performance are achieved.