117 results about "Supersonic speed" patented technology

The invention belongs to the field of target drone fuel oil systems, and particularly relates to a self-adaptive supersonic speed target drone fuel oil system and a control method thereof. The system comprises a machine body oil tank, a first electromagnetic valve, an overflow valve, a pressurization oil tank, a fuel pump, an engine pump, a monitoring module and a control module. The body fuel tank comprises a main fuel tank and an auxiliary fuel tank which are arranged in a fuel tank in the middle of the target drone body and located in front of and behind the theoretical mass center of the target drone respectively. The pressurization oil tank is additionally arranged in the fuel oil cabin, fuel oil in the engine body oil tank enters the pressurization oil tank for pressure maintaining after being pressurized by the fuel oil pump, and then oil is supplied to the engine from the pressurization oil tank, so that stable oil supply to the engine is achieved by establishing a local pressure environment instead of a large pressure environment of the whole fuel oil cabin, and therefore when a fuel oil system works, the fuel oil can be stably supplied to the engine. The oil supply flow and pressure are controlled in a closed-loop mode in real time through the control system according to the engine state.

The invention discloses an anti-oxidation and anti-scouring heat prevention and insulation composite material, a preparation method and application, and belongs to the technical field of high-temperature thermal protection materials. According to the method, a two-step in-situ modification method is adopted, hydrolyzed organosilane reacts with phenolic resin to prepare silicon modified phenolic resin, then the silicon modified phenolic resin is bonded with an organoboron compound to obtain boron-silicon co-modified phenolic resin, a curing agent is added, then the boron-silicon co-modified phenolic resin is injected into a fiber reinforcement mold, and a finished product is obtained through a sol-gel reaction and normal-pressure drying. According to the invention, boron and silicon elements are embedded into phenolic aldehyde molecular chains through covalent bonds to form a boron-silicon-carbon hybrid network, a nano ceramic phase and a compact glassy protective layer are generated at high temperature, and the coating has excellent anti-scouring property, thermal stability and heat insulation property. The preparation process is simple and controllable, is suitable for large-scale production, can accurately regulate and control the performance, and is suitable for thermal protection parts of supersonic aircrafts.

The invention discloses a rotating detonation afterburner based on centrifugal nozzle discrete injection, and belongs to the technical field of detonation propulsion. The combustion chamber comprises a gas collecting cavity, an injection unit, a centrifugal nozzle, an outer ring and an inner column, wherein an annular detonation combustion chamber is formed between the inner wall surface of the outer ring and the outer wall surface of the inner column. The centrifugal nozzles are uniformly arranged at the intersection of the gas collection cavity and the injection unit in the circumferential direction to realize primary atomization of the fuel oil; the air is accelerated by the injection unit from the air collection cavity and then is mixed with the atomized fuel oil, secondary atomization is achieved, the particle size of liquid drops is remarkably reduced, and the fuel activity and mixing uniformity are improved. The injection unit adopts a Laval nozzle profile, air is accelerated to supersonic speed in the injection unit, and pressure of a combustion chamber is prevented from being transmitted forwards. The rotary detonation engine is suitable for the high-temperature low-oxygen environment, supports the annular combustion chamber configuration and the empty cylinder combustion chamber configuration, can achieve stable rotary detonation detonation and propagation of liquid fuel and air, and provides technical support for engineering application of the rotary detonation engine.

The present disclosure is directed to a supersonic oblique rotating detonation wave engine (SORDE) and systems and methods for generating a supersonic oblique rotating detonation wave. The SORDE is configured to produce and sustain a supersonic oblique rotating detonation wave through the injection of fuel at supersonic speeds into an inlet air flow between Mach 1 and Mach 7. The SORDE and method include injecting fuel into the inlet air in an amount to generate an equivalence ratio of 0.2 to 2.5. Some embodiments include a plurality of fuel injector ports each having a diameter of about 0.010 inches to about 0.040 inches; an annular wedge disposed in or upstream of the detonation chamber with an angle of about 5 degrees to about 40 degrees relative to a longitudinal axis of the engine; and / or a cylindrical center body disposed in or upstream of the detonation chamber.

The application provides a rotary impingement oblique flow centrifugal combined compressor and application thereof, and belongs to the technical field of turbine engines, and comprises an outer casing structure, an inner casing structure, a working mechanism and a diffuser mechanism, the outer casing structure and the inner casing structure are rotary bodies, the inner casing structure is located inside the outer casing structure, the inner casing structure and the outer casing structure form a working cavity, the working mechanism is installed in the working cavity and is used for pressurizing airflow. Compared with the compression mode of a conventional combined compressor, the application draws lessons from the pressurization mode of a supersonic inlet, and the Mach number of airflow relative to a rotating impeller reaches 2 or more through high-speed rotation; at this time, a series of weak oblique shock waves are constructed through the contraction of a passage, and a high-efficiency compression system is obtained through the pressurization of the shock wave system, the single-stage pressure ratio of the high-efficiency compression system is much higher than that of a conventional compressor, and the application has the advantages of compact structure, simple processing and long service life.

The invention relates to the field of metal nano material preparation, and discloses a supersonic gas atomization preparation method of high-dispersity spherical nano aluminum powder, which comprises the following steps: establishing a constant-temperature aluminum melt flow, and constructing a supersonic gas-solid two-phase atomization jet flow; nitrogen with the temperature lower than the melting point of the particles is used for carrying flow, solid zinc stearate particles are carried and accelerated to supersonic speed, and the aluminum melt flow is impacted; solid rigid momentum of the particles penetrates through an aluminum liquid interface to enter liquid drops, thermal relaxation time difference is used for inducing the particles to generate phase change decomposition in the liquid drops, the liquid drops are broken from the inside through volume expansion, and a carbon-zinc composite isolation layer is formed in situ through decomposition products. And high-efficiency nanocrystallization crushing under low energy consumption is realized.

The present application belongs to the technical field of hypersonic wind tunnel test, and discloses a test checking method of hypersonic frictional resistance standard value. The test checking method of hypersonic frictional resistance standard value is carried out in a hypersonic quiet wind tunnel, and the model is stably kept in a laminar flow state by means of a quiet flow field provided by the hypersonic quiet wind tunnel, so that the interference of flow field noise on the hypersonic frictional resistance is eliminated, real near space flight test data is obtained, and the obtained hypersonic frictional resistance standard value can be used for related test calculation or theoretical analysis, thereby providing technical support for the research on the frictional resistance of a spacecraft.

The invention provides a rudder for a supersonic aircraft and a manufacturing method, and relates to the field of supersonic aircraft design, the rudder is integrally formed through co-curing, the rudder comprises a rudder body, the rudder body comprises a supporting plate, the supporting plate is formed by co-curing a carbon fiber composite material, and the supporting plate is made of a carbon fiber composite material; the supporting plate is provided with at least three hollow supporting pipes, and the cross section of each supporting pipe is in a regular hexagon shape. The supporting inner core is arranged in the hollow position of the interior of the supporting pipe, and the cross section of the supporting inner core is in a regular hexagon shape; the supporting outer layer covers the surface of the supporting plate; a heat protection layer; the supporting inner core and the supporting outer layer are made of foam materials. According to the rudder for the supersonic aircraft and the manufacturing method, the overall structure of the rudder is co-cured and integrally formed, the supporting plate made of the carbon fiber composite material is adopted, the stability of the structure of the rudder is guaranteed, meanwhile, the overall mass of the rudder is effectively reduced, and meanwhile mass production is facilitated.

The present application relates to the technical field of rail production process, in particular to a production method for improving wear resistance of channel rail for tram. After universal rolling, compressed air is applied to the channel rail for forced cooling in parts. In the first stage, supersonic air jet cooling is applied to the rail head tread, rail head running edge, rail head left side, rail lip and rail bottom to control the cooling speed of each part. In the second stage, supersonic air jet cooling is continuously applied to each part of the rail to control the cooling speed of each part. In the third stage, subsonic air jet cooling is applied to each part of the rail to control the cooling speed of each part, and then the rail is naturally cooled in air. The quenching effect of the channel rail tread and running edge is improved, the pearlite interlamellar spacing is refined, the rail strength and hardness are improved, and the channel rail has good wear resistance.

A system for capturing a target condensable gas may include: a supersonic condensation nozzle configured to receive a gas flow having a target condensable gas content, and for accelerating said gas into a supersonic gas flow, at least one particle injector in the supersonic nozzle for injecting particles for the particles to be in the supersonic gas flow, whereby target gas condensate forms onto some of the particles, an expansion device defining a conduit having a convex corner causing the flow of the particles with target gas condensate to generate expansion waves to separate the flow into a first flow path having an increased concentration of particles with target gas condensate thereon and into a second flow path of gas with lowered target condensable gas content, and a divider having at least a first conduit, the divider configured to collect the first flow path.

The invention relates to a hypersonic speed temperature sensor measuring device, which relates to the technical field of temperature measurement, and comprises an outer tube, a stagnation cover and an exposed end type sheathed thermocouple, one end of the outer pipe is closed and the other end is open; a plurality of stagnation covers communicated with the outer pipe are formed on the outer side wall of the outer pipe in the length direction of the outer pipe, air inlets are formed in the ends, away from the outer pipe, of the stagnation covers, and at least one air outlet is formed in the side wall of each stagnation cover; an exposed end type sheathed thermocouple is arranged in each stagnation cover, and a metal shell of the exposed end type sheathed thermocouple is hermetically welded with the stagnation cover; and all the air outlets are positioned between the detection ends of the corresponding exposed end type sheathed thermocouples and the metal shell. The hypersonic speed temperature sensor measuring device is reliable in overall structure, the temperature collecting duration of the hypersonic speed temperature sensor measuring device under high-temperature hypersonic speed airflow impact can be prolonged, and the credibility of a measuring result is improved.

A supersonic diffuser for use in turbomachinery adapted to impart thermal energy to a process fluid is provided. The diffuser includes a vaned zone configured to define a passageway having a flow area to pass a flow of a process fluid at supersonic velocity. The vaned zone is configured to define a step-change to the flow area of the passageway at a given location, such as where the flow of the process fluid exits the vaned zone. The diffuser further includes a shock zone coupled to the vaned zone to pass the flow of the process fluid that exits the vaned zone. The shock zone is configured to support a system of shock waves that increases static temperature of the process fluid downstream of the system of shock waves. A mixing and subsonic diffusion is configured to decelerate process fluid from the shock zone to a reduced subsonic speed prior to discharge of the process fluid through an exit of the diffuser.

The invention discloses an oil-stain-resistant and wear-resistant efficient protective rare earth coating and a preparation method thereof. The coating is composed of 50-70 parts of tungsten carbide powder, 10-20 parts of cobalt-coated tungsten carbide composite powder, 8-15 parts of yttrium-stabilized zirconia powder, 5-12 parts of rare earth oxide composite powder and 5-10 parts of nickel-chromium-aluminum-yttrium alloy bonding powder, wherein rare earth oxide is a mixture of cerium oxide and lanthanum oxide. The preparation method of the coating comprises the steps of base body sand blasting coarsening and cleaning, three-dimensional mixing of all component powder and high velocity oxy-fuel spraying deposition. Decarburization of tungsten carbide is inhibited through a cobalt coating structure, composite rare earth is utilized to refine grains and reduce surface energy, and an optimized spraying process is combined, so that the coating has high hardness, high bonding strength, low porosity, low wear rate and excellent oleophobicity, the problems that a traditional wear-resistant coating is easy to adhere, difficult to clean and short in service life in a greasy dirt environment are effectively solved, and the service life of the coating is prolonged. The coating is suitable for surface protection of key parts such as hydraulic machinery and aero-engines.

A supersonic aircraft is disclosed herein. The supersonic aircraft comprises one or more fuselages, one or more wings, at least one wing-mounted engine, at least one inlet mounted on the upper surface of or above the wing, an upper wing surface propulsion system fairing, and a lower wing surface propulsion system fairing or lower wing surface modification to account for the presence of the engine.

An energy system converts the latent heat energy of moist air into water and in some examples also to motive energy in dry exhaust air. It does this by flash evaporation (5) of primary air which has added water droplets (6), the supersonic speed of the primary flow from the flash evaporation nozzle creating suction to draw in ambient secondary air. A condenser tube (10) with a mixing throat (12) enhances the mixed primary / secondary air speed and causes condensation within the volume of the tube without need for heat exchanger surfaces, to provide output water and dry air at a pressure higher than atmospheric and thus providing thrust / power.

The invention relates to mufflers (1) that suppress the sound of a bullet (5) fired from a firearm, preferably at supersonic speed, and maintain the accuracy and distance of the shot. The weapon muffler (1) includes a body (2) mounted on its barrel (3) by means of a bracket (4), which has a chamber (7) for expanding gases accompanying the bullet (5) and a bullet passage channel (8), with a diameter of 1.1+3 times its caliber and at least 1.5 times this diameter, at least in the form of a cylindrical and / or conical hole (8), which has grooves or thread (10) on its inner surface, or these are formed by a metal mesh tube inserted into it.

The application relates to the technical field of hypersonic aircrafts, and particularly discloses a heat protection and power generation integrated system and method based on a semi-open Rankine cycle, which comprises a water storage tank, a hydrogen storage tank, an electrolytic cell, a steam turbine, a heat exchanger, a hydrogen turbine and a combustion chamber. The system uses liquid water to absorb heat at the leading edge of the aircraft, vaporizes into water vapor to drive the steam turbine to generate power, realizes active heat protection and waste heat conversion; the water vapor after work exchanges heat with liquid hydrogen in the heat exchanger to condense, preheats the fuel while completing the Rankine cycle; the vaporized hydrogen drives the hydrogen turbine to generate power again, improving the power generation efficiency; the condensed water can be introduced into the electrolytic cell to be electrolyzed into hydrogen and oxygen, and the generated hydrogen directly supplements the fuel. The system realizes efficient heat protection, continuous power generation, in-situ production and supplement of fuel through structure coupling, and effectively reduces the mass of the homeward journey system by converting water into fuel in the middle and later stages of flight, thereby improving the endurance and high-altitude performance of the aircraft.

The utility model discloses a sonic torch burner. The sonic torch burner comprises a Laval nozzle, a gas inlet of the Laval nozzle is connected with a gas outlet of the torch gas pipe, and a pressure regulating valve is arranged at one end, far away from the Laval nozzle, of the torch gas pipe; a flow velocity sensor for monitoring the airflow velocity is arranged at the narrow throat of the Laval nozzle; and the pressure regulating valve and the flow velocity sensor are connected with the controller. The utility model further discloses an elevated torch and a ground torch. Through the combined control of the pressure regulating valve, the flow velocity sensor and the controller, the airflow velocity at the narrow throat of the Laval nozzle is ensured to be sound velocity, so that the stable supersonic speed regulation and control of the combustible gas velocity at the outlet of the Laval nozzle are ensured, and the combustion stability is further realized.

The invention discloses a carbon fiber reinforced (TiZrHfNbTa) C ceramic-based composite material as well as an interface regulation and control preparation method and application thereof, and belongs to the technical field of ultrahigh-temperature ceramic-based composite materials. The carbon fiber reinforced (TiZrHfNbTa) C ceramic-based composite material comprises carbon fibers and a ceramic matrix, a graphene bonding layer interface and a pyrolytic carbon layer interface are arranged between the carbon fibers and the ceramic matrix, the pyrolytic carbon layer interface and the graphene bonding layer interface sequentially wrap the carbon fibers from inside to outside, the interface bonding strength is high, the toughening effect of the fibers is fully exerted, and the carbon fiber reinforced (TiZrHfNbTa) C ceramic-based composite material is suitable for being used as a high-temperature-resistant composite material. The mechanical property of the composite material is remarkably improved, meanwhile, the composite material has the high-temperature-resistant characteristic of ultra-high-temperature ceramics, and the composite material has wide application prospects in extreme environments such as thermal protection systems in the aerospace field and leading edge components of hypersonic aircrafts.

The invention discloses a method for preparing a ceramic-based composite material through ultra-fast infiltration and pyrolysis and application, and belongs to the field of ceramic-based composite material preparation. The preparation method comprises the following steps: dipping a porous C / C composite material blank or a fiber preform in a ceramic precursor, carrying out cross-linking curing at a certain temperature, and then under a vacuum or inert atmosphere, adopting a Joule heating method to make the porous C / C composite material blank or the fiber preform subjected to ultra-fast heating (gt; and carrying out heat preservation at the temperature of 10 DEG C / s to finish a precursor-ceramic phase conversion process, and then repeating impregnation, cross-linking curing and precursor-ceramic phase conversion to realize the ultra-fast preparation of the ceramic-based composite material. The method can effectively solve the problems of long period, high energy consumption and the like of an impregnation cracking process. The method can be applied to ultra-fast preparation of high-temperature components such as hypersonic aircraft nose cones, wing leading edges and engine exhaust nozzles, and the mechanical property and the heat-proof performance of the high-temperature components can be improved.

In order to solve the problems that in the prior art, calculation precision and efficiency are difficult to consider at the same time during supersonic jet machining, and detached shock waves cannot be accurately captured, the invention provides a finite element simulation method for the stagnation flexible constraint flow field characteristics of a jet impact area. The finite element simulation method comprises the steps that firstly, a geometric model of a jet area and an outer watershed is established, and preliminary solving is conducted; extracting jet flow axis pressure gradient characteristics to determine detached shock wave interval coordinates; then reconstructing a block geometric model comprising an upper jet flow region, a detached shock wave region, a stagnation region and an outer watershed according to the coordinates, and implementing differential local grid encryption on the shock wave region and the stagnation region; and finally, carrying out second iterative solution based on the Sutherland viscosity law to obtain final flow field distribution. According to the method, the thickness and the position of the shock wave layer are accurately captured while the number of the grids is controlled through the block encryption strategy, the contradiction between the calculation economy and the numerical precision under the transonic flow condition is solved, and the method is suitable for numerical simulation of the complex supersonic gas jet flow field containing the detached shock wave and the stagnation region.

This invention discloses a shockwave cyclone reactor and method for biomass flash pyrolysis. The reactor is a tubular structure and includes at least: a shockwave preheating section, which includes a tubular shell and a rotating shaft disposed within the shell, with a circumferential gap between the rotating shaft and the tubular shell forming a gas channel; the gas channel is provided sequentially along the gas flow direction by: a first stator for initial acceleration of the process gas, a rotor for accelerating the process gas to supersonic speed, and a second stator for generating shockwave losses and rapidly raising the temperature of the process gas; a biomass particle inlet is located downstream of the shockwave preheating section and arranged tangentially along the circumference of the tubular reactor, for mixing the biomass particles with the shockwave-heated gas; guide vanes for forming a cyclone are provided between the biomass particle inlet and the shockwave preheating section; a pyrolysis section is located below the biomass particle inlet and includes a cylindrical section and a conical section; the biomass particles and high-temperature gas form a cyclone in the cylindrical section, pyrolysis is performed in the cyclone state, and gas-solid separation is performed in the conical section.

The invention relates to the technical field of wind-water fire extinguishers, and discloses an ultra-high-speed vortex jet wind-water fire extinguisher with double supports and uniform load distribution, which solves the problems in the prior art. The device comprises an upper shell, a lower shell, a control bin and an injection mechanism, the upper shell and the lower shell are fixedly connected through bolts; the spraying mechanism comprises a pre-atomization barrel fixedly connected to the inner wall of the lower shell through a fixing shaft, the inner wall of the lower shell is fixedly connected with a Laval air pipe, the Laval air pipe is arranged on the side wall of the pre-atomization barrel in a sleeving mode, and through a three-stage atomization system composed of the pre-atomization barrel, the Laval air pipe and the mixing barrel, the spraying mechanism can spray liquid in the mixing barrel. Micro-scale shearing, vortex collision and secondary crushing of supersonic airflow to a rotating water film are achieved inside, a micron-scale aerosol-state fire extinguishing medium is generated, the fire extinguishing medium is high in penetrating power, capable of bypassing obstacles to extinguish deep fire and blind fire, long in suspension time and capable of continuously dipping a fire scene, after-combustion is fundamentally prevented, and the fire extinguishing efficiency and effect are far better than those of a traditional technology.

The present disclosure provides a nozzle design that allows compression and / or condensation of gas or vapors presence in a liquid entering the nozzle. In the nozzle configuration of the present disclosure, the stream that flows through the nozzle is a two-phase steam and it reaches a supersonic flow before the suction of the additional vapors begins. By the end of the first narrowing section, the flow within the nozzle is supersonic, followed by a first expanding section that further reduces the pressure of the stream and increases the Mach number before the suction section that follows it, which maintains the pressure and the supersonic conditions.

The application relates to a staged pre-cooling single-wall nozzle extension section, which comprises a single-wall nozzle extension section upper section and a single-wall nozzle extension section lower section, turbine exhaust gas is proportionally divided into two streams, no more than 30% of the turbine exhaust gas forms a first supersonic speed gas film, a lower-temperature protective layer is formed in a region close to the inner wall surface of the single-wall upper section base body, so as to separate the higher-temperature main flow gas and the single-wall upper section base body, a second supersonic speed gas film is arranged at a position where the first gas film cooling efficiency is reduced to 80% or the maximum wall surface temperature of the single-wall nozzle extension section upper section approaches the material allowable temperature, and the remaining inner wall of the single-wall section is heat-protected. The application significantly reduces the maximum temperature of the structural wall surface under specific flow distribution.

The application provides a target projectile and an aircraft. The target projectile comprises a head part in a conical shape, a projectile body comprising at least one engine body section, a first engine body section close to the head part being connected to the head part, a tail part arranged at an end of the engine body section away from the head part and comprising at least two tail wings distributed around an axis of the engine body section, and a rudder part arranged at an end of the first engine body section close to the head part and comprising at least two rudder wings distributed around an axis of the first engine body section, wherein in a first plane perpendicular to the axis of the engine body section, the orthographic projection of the at least two rudder wings is staggered with the orthographic projection of the at least two tail wings. The application realizes controllable supersonic flight of the target projectile in a large airspace and a wide speed range, with a flight height of 0.1-20 km and a flight speed of 0.3-2.4 Ma.

Multi-layer protective coating systems, methods of making such coatings, and articles of manufacture provided with such coatings. Such a coating system has a SiC layer overlying a surface of a carbon-fiber reinforced carbon matrix (C / C) composite, a ZrB2—SiC layer overlying the SiC layer, a ZrC—ZrO2 layer overlying the ZrB2—SiC layer, and a Sm2O3—ZrO2 layer overlying the ZrB2—SiC layer and preferably serving as a topcoat of the coating system. The coating system is capable of serving as an ablation and oxidation protection measure for C / C composites, for example, as may be experienced at the leading edge(s) of a hypersonic aeronautical vehicle while moving through the earth's atmosphere.

The present disclosure provides a fin stabilized projectile which can accommodate subsonic and supersonic flight velocities, wherein flight performance characteristics of the projectile are stabilized. The projectile is configured to be incorporated into a sabot free cartridge, wherein the projectile can be launched from a smooth bore barrel. The projectile includes an elongate body having a diverging ogival nose section extending rearward from a leading end, a converging ogival back section terminating at a trailing end; and a tail segment having a tail fin projecting radially from the back section and extending longitudinally forward of the trailing end of the elongate body and extending longitudinally rearward of the trailing end of the elongate body.

The supersonic air inlet channel comprises an outer pressure section, an inner pressure section, an expansion section, the lip mask, a high-pressure air source and a channel, wherein the lip mask is located on the outer side of the inner pressure section and provided with the locally-deformable air bag, and the channel is connected with the high-pressure air source and a middle cavity of the air bag. The air inlet channel is a supersonic air inlet channel, and the working range is Ma0-3 +. Deformable air bags are arranged on the front edge and the downstream of the lip mask, a lip mask body is of a rigid structure with a cavity in the middle, deformable air bag installation grooves are formed in the two sides of the lip mask body, the walls of the installation grooves are porous walls, and connection between the cavity and the deformable air bags is achieved. The pressure of the cavity is adjusted through the high-pressure air source and the deflation valve, and the configuration of the deformable air bag is changed. The cavity pressure of the air inlet channel is reduced from the subsonic-speed working condition to the supersonic-speed working condition, the front edge of the lip cover is transited from the blunt front edge to the sharp front edge, airflow separation on the inner side of the lip cover under the subsonic-speed working condition is restrained, the shock wave strength of the front edge of the lip cover under the supersonic-speed working condition is weakened, the performance of the air inlet channel is improved, the overflow resistance of the air inlet channel is reduced, and the installation thrust of an engine is increased.

The invention discloses a speed measurement system and method suitable for a combined pulse high-enthalpy wind tunnel, and belongs to the technical field of aerodynamic tests. The objective of the invention is to improve the speed measurement precision in a high-speed high-enthalpy unsteady flow field. The system comprises a water cooling machine, a femtosecond laser, a reflector I, a time schedule controller, an input power supply, a reflector II, a reflector III, a plano-convex lens, a wind tunnel test section, a computer device and an imaging system. The femtosecond laser is respectively connected with the water cooling machine, the time sequence controller and the input power supply, and the imaging system is respectively connected with the time sequence controller and the computer device; laser emitted by the femtosecond laser irradiates the reflector I to be reflected, then is reflected by the reflector II and the reflector III and then enters the plano-convex lens, emergent laser enters a wind tunnel test section, and the imaging system shoots a test photo of the wind tunnel test section. Effective support is provided for hypersonic pneumatic tests and aircraft design verification.