88 results about "Expansion wave" patented technology

Devices are provided for regulating the pressure pulses and flow pulses that are emitted in the exhausts of apparatus and equipment such as internal combustion engines, compressors and poppet valves. The pressure pulses are reflected from a reflecting member to their source as expansion waves that create a reduced pressure zone that allows subsequently emitted flow pulses to be more easily emitted, thereby enhancing the operational efficiency of the apparatus and equipment. Flow pulses that contain exhaust products such as the products of combustion from an internal combustion engine are regulated so that the exhaust flow they generate is not unduly restricted within the device and an undesirable back pressure that resists the emission of flow pulses from their source created. In regulating the pressure and flow pulses, the devices attenuate the noise generated by the pressure and flow pulses.

The invention provides a method for calibrating a supersonic flow field density field. According to the method, a supersonic flow field density-nanoparticle-based planar laser scattering (NPLS) image gray curve is calibrated by adopting a comprehensive oblique shock wave and expansion wave calibration method based on an NPLS technology. The method comprises the following steps of: 1, uniformly scattering trace particles in supersonic incoming flow, and shooting a particle image according to the instruction of a computer by a charge coupled device (CCD); 2, continuously adjusting obliqueness of an attack angle alpha in a supersonic wind tunnel, and acquiring a group of (rhoi, Ii) (i=1, 2, L, n-1) data by changing the oblique attack angle alpha; 3, placing expansion wave generators in the supersonic wind tunnel, and acquiring the other group of (rhoi, Ii) (i=n, n+1, L, N) data by placing the expansion wave generators with different deflection angles; and 4, performing polynomial fitting on the two groups of data to obtain the supersonic flow field density-NPLS image gray relation curve, namely rho=alapha0+alpha1I+alpha2I2+alpha3I3+K. The method aims to solve the technical problems of low spatial resolution and low signal-noise ratio and high error on measurement of a low-density area.

The invention relates to a jet ignition device of a pulse detonation engine, which is characterized in that a jet tube is fixedly connected with a mixing chamber through a sealing clamp nut, an adjusting plug is mounted at the top end of the jet tube through a sealing nut and a clamping nut; a drain tube is arranged in a jet chamber inside the jet tube and is connected with the jet tube through a fixing rib, and the lower end of the drain tube extends into the mixing chamber; and an ignition seat is arranged on one side of the jet tube and is located at one end of an outlet of the drain tube. Compared with the time and the distance for generating detonation waves by normal ignition devices, the time and the distance for generating the detonation waves by the jet ignition device are shorter, accordingly the DDT (deflagration to detonation transition) time and distance is shortened. In this case, length of a detonation chamber can be further shortened so as to improve working frequency of the pulse detonation engine. After the detonation chamber is shortened, exhausting burden in the pulse detonation engine is reduced, so that expansion wave-beams generated from the outlet of the pulse detonation engine can catch up with pressure waves transmitted before, pressure disturbance in an air-inlet passage is reduced, and working stability of the pulse detonation engine is enhanced while better propulsive performance is achieved.

The invention provides an ultrasonic nonuniform flow nozzle and a design method thereof. The design method for the ultrasonic nonuniform flow nozzle comprises the following steps of: determining a subsonic speed section curve according to the diameter of an inlet of a subsonic speed section; determining a left characteristic line boundary and a right characteristic line boundary of an outlet dependence domain according to Mach number distribution of an outlet and a method of characteristics; determining an initial characteristic line from a throat; determining an equi-flow line of the nonuniform flow nozzle; and determining upper and lower wall curves by using the initial characteristic line and the equi-flow line according to conservation of mass and the method of characteristics. By thedesign method for the ultrasonic nonuniform flow nozzle, the ultrasonic nonuniform flow nozzle can be designed, so that the ultrasonic nonuniform flow nozzle generates the required nonuniform flow, and can continuously eliminate waves, eliminate or reduce centralized compression or expansion waves and ensure the quality of a flow field. According to the ultrasonic nonuniform flow nozzle, the design method for the ultrasonic nonuniform flow nozzle is used.

The invention discloses a method and a device for implementation of a supersonic runner. The method includes: acquiring a first center point on the center line of the runner according to constraint parameters of an entrance flow field; acquiring a second center point on the center line of the runner according to constraint parameters of an exit flow field; and generating a wall face curve of the supersonic runner according to a third center line on the ligature of the first center line and the second center line, preset wall face points of a circular runner and preset wall face points of a rectangular runner. The problem that the single compression characteristic of a single standard flow field in correlation technique cannot simultaneously guarantee match of expansion waves and compression waves in the rectangular pipeline and the circular pipeline, and the design of the supersonic runner with inner runner section changed from circular to rectangular is more flexible.

The invention discloses a laminated plate aircraft skin with lift augmentation, resistance reduction, flow control and deicing prevention functions. The laminated plate aircraft skin comprises a glass fiber layer (2) and an aluminum alloy (1), wherein the glass fiber layer (2) and the aluminum alloy (1) are alternative; the surface of the skin is provided with an insulating material layer (4) and an electrode layer; the electrode layer is divided into an upper layer and a lower layer; the upper layer and the lower layer are respectively positioned on two sides of the insulating material layer (4) and are asymmetrically distributed; the upper electrode layer (3) is exposed in air and is connected with an alternating current power supply; and the lower electrode layer (5) is positioned below the insulating material layer (4) and is grounded. The laminated plate aircraft skin can be directly applied to aircraft manufacturing engineering, cannot be damaged by high-speed airflow and cannot bring parasite drag due to the damage of the original pneumatic shape. The deicing prevention function is realized by utilizing plasma heating on air and the pulse expansion wave function, the flow separation can be effectively controlled, the stall is delayed, the highest lift coefficient is increased, and the pressure difference resistance is reduced.

The present invention relates to an ultrasonic blade and a design method applied to aeroengine fans or aerocompressor rotors, pertaining to the art of impeller mechanical technology. The blade is characterized in that the small part of the blade on the front edge is S-shape and the large rear at the rear is flat and the thickness of the blade gradually increased from front to back; the ultrasonic air current flows into a grid formed by the blade, produces an expansion wave in the entrance of the grid, consequently forms an oblique shock wave in the transition from the S-shaped suction surface of the blade to the flat section intersected with the front edge of the adjacent blade, then produces a reflection shock wave and reproduces a local shock wave near the exit of the grid. The design method of the blade includes the following process: the front part is confirmed to be S-shaped and the rear part is divided by a straight medial camber line; the thickness distribution of the blade, which gradually increases from front to back, is confirmed; the surface mold of the blade is confirmed according to the medial camber line and the thickness distribution of the blade. The blade provided by the present invention can effectively reduce the flowing loss, increase the rim power of the rotor and improve the overall pressure ratio.

The invention discloses a modified scramjet combustion chamber and a design method of a swirler thereof. The swirler capable of causing air flow to generate helical motion is arranged in a combustion chamber and comprises a rotational flow generation upper wallboard, a rotational flow generation lower wallboard, a back end surface wallboard and a sidewall mounting surface, wherein the rotational flow generation upper wallboard is used for making the air flow entering the combustion chamber generate oblique shock waves, and the rotational flow generation lower wallboard is used for making the air flow entering the combustion chamber generate expansion waves. The sidewall mounting surface is connected with the combustion chamber, and the front edge of the swirler faces to an inlet of the combustion chamber; the back end surface wallboard is arranged opposite to an outlet of the combustion chamber; the interior of the swirler is provided with a fuel conveying embedded pipe; an inlet of the fuel conveying embedded pipe is arranged on the sidewall mounting surface; and an injection opening of the fuel conveying embedded pipe is arranged at the back end surface wallboard. Therefore, the invention has the advantages of improving the mixing efficiency and the combustion efficiency of the fuel of a scramjet, shortening the lengths of an engine isolating section and the combustion chamber, further optimizing the structure of the engine, increasing the thrust-weight ratio of the engine and improving the whole performances of the engine.

The invention discloses a chemical flooding alternate injection method, relates to optimization of a chemical flooding injection mode, and solves the problems of imbalanced distribution of injection rate, greatly reduced effective utilization ratio and low low-permeability oil layer producing degree simultaneously caused by the conventional ternary compound system general injection. The method comprises the following steps of: 1, injecting polymer plug A into the pore until the polymer plug A is in amount which is 10 percent of pore volume; 2, injecting ternary plug into the pore until the ternary plug is in an amount which is 10 percent of pore volume; 3, repeating the step 1 and the step 2 for 3 to 4 times; and 4, injecting polymer plug B into the pre until the polymer plug B is in an amount which is 10 to 12.5 percent of pore volume to finish chemical flooding injection. Compared with the general injection mode, the chemical flooding injection method of the invention has the advantages that: compound system expansion wave and volume effect are enhanced, the liquid absorption of a chemical agent by a low-permeability layer is improved, the effective utilization ratio of the chemical agent is improved and the yield is further improved by about 5 percent indoors under the condition of the same using amount.

The invention discloses a pressure wave supercharging auxiliary twin-stage vapor compression refrigeration system and a working method thereof and belongs to the technical field of mechanical refrigeration. The system utilizes the supercharging characteristic of a flashing shock wave to complete precompression of refrigerant vapor and realize the purpose of reducing the pressure ratio of a refrigeration compressor; two-phase separation of a liquid phase refrigerant is completed by utilizing the expanding characteristic of an expansion wave to realize twin-stage refrigerating cycle of isentropic expansion and isenthalpic throttling. A refrigerant enters a four-port wave rotor supercharger via an evaporator to be supercharged and mixed with airflow of the gaseous phase opening of a separator, enters the refrigeration compressor to be supercharged, and is condensed to enter the four-port wave rotor supercharger in a saturation liquid manner for flashing; formed two-phase mixture passes through the separator, the liquid phase throttles to enter the evaporator for refrigeration, the gaseous phase is mixed with supercharged vapor, and the mixture enters the refrigeration compressor. The pressure wave supercharging auxiliary twin-stage vapor compression refrigeration system is simple, solves the problems that the compressor is large in pressure ratio and has no stage intercooling of multistage compression during the vapor compression refrigeration cycle, and has the advantage of being higher in energy efficiency ratio than the traditional twin-stage vapor compression refrigeration system.

The invention provides a hedge damping gas wave refrigerator which belongs to the compressed air expansion refrigeration technical field. The device can realize refrigeration through the moving shock wave and the expansion wave produced in the oscillation tube by the air pressure. The hedge damping gas wave refrigerator is characterized in that the radiation tubes or the oscillation tubes arranged in a circular shape are arrayed in pairs with the same degree of angle between each other, and a damping dissipative structure is connected with the ends of each pair to put through the flow of the two pipes, thus reducing the rigidity of the air compression of the oscillation tubes; thereby the device can constantly run with high speed in working conditions that the short pipe is too long and the rotation rate is low. The device of the invention has the advantages that the structure is simple, the assembly and disassembly are convenient, the condensation and stagnancy of the liquid is avoided, the adaptability to variable working conditions is high, the isentropic efficiency of refrigeration is high, etc. The device of the invention can produce low temperature and can be widely used in the technical fields such as the liquefied separation of mixed gases, the recovery of light hydrocarbon in petroleum gases, the liquefaction of natural gas, and the supply of low-temperature air-conditioning source, etc.

Cold-heat effect of shock oscillation tube combined with jet oscillation technique is adopted in the invention to reach purpose of refrigeration. The machine is composed of main machine body, static ejector nozzle, oscillation receiving tube and acoustic oscillator etc. Under combined action of gas distributor and acoustic oscillator, gas is jetted in periodic deflexion and periodic intermittent into receiving tubes around to generate flow in high frequency oscillation. Further, under interaction effect and inference of series of wave including shock wave, compression wave and expansion wave, the machine completes unsteady expanding work and energy conversion so as to realize refrigeration through gas. Features are simple structure adjustable continuous flow, suitable to work medium in phases of gas and liquid.

The invention relates to a method for calculating micro-pressure waves generated in the process of passing through a tunnel by a high-speed train. The method for calculating micro-pressure waves is characterized by comprising the following steps: dividing a flow field of the tunnel into a plurality of areas by using a dynamic mesh; calculating areas containing a first compression wave and an expansion wave and an area before the compression wave and the expansion wave; and simulating to obtain a micro-pressure wave value of a tunnel outlet. The method for calculating the micro-pressure waves generated in the process of passing through a tunnel by a high-speed train better solves the problems of oversize scale, overlong calculating period, and the like of a three-dimensional calculating mesh of a long tunnel and a tunnel group. In the method, the flow field is divided into a plurality of the areas by adopting dynamic mesh technology and the value simulation on the micro-pressure waves of the tunnel outlet are carried out by only calculating the areas containing the first compression wave and the expansion wave and the area before the first compression wave and the expansion wave according to the forming mechanism of the micro-pressure waves. Meanwhile, the invention provides a new method for the calculation of the micro-pressure waves of the tunnel outlet.

An ejector includes a nozzle having a fluid passage circular in cross section. The fluid passage includes a throat portion smallest in cross-sectional area, a divergent portion that becomes larger in cross-sectional area toward a downstream side from the throat portion, and an ejection port that is provided at a downstream end of the divergent portion. A passage wall surface of the divergent portion includes a recess portion that is recessed from within outward in a radial direction of the passage wall surface. The recess portion is located adjacent to the ejection port, and the recess portion extends continuously in a circumferential direction of the passage wall surface to enclose the fluid passage and have an annular shape. Accordingly, noise due to an expansion wave of the ejected fluid can be reduced.

The invention provides a Mach number sensitivity analysis method for supersonic cavity flow, and the method comprises the steps: forming expansion waves at different angles at the front edge of a flat plate through adjusting the attack angle of a model, wherein the Mach number will rise after a gas flow passes through the expansion waves; and carrying out the sensitivity analysis of the Mach number through the comparison of the pulsating pressure data before and after the change of the attack angle. The method can achieve the Mach number sensitivity analysis under the supersonic conditions, and solves a key problem of a wind tunnel sensitivity analysis test.

The invention relates to an on-line upgrading capacity expansion power adjusting method of a wavelength division multiplexing optical transmission system. The method comprises the steps of: confirming the transmitting power requirement difference between an upgrading capacity expansion wave channel and the original wave channel; taking the original wave channel observed at a circuit monitoring hole as reference; and on the basis, adjusting the power of the upgrading capacity expansion wave channel, observing from the circuit monitoring hole, and leading the difference between the upgrading capacity expansion wave channel and the original wave channel to reach the requirement. At the moment, the transmitting optical power of the capacity expansion wave channel on the circuit reaches the requirement, OSNR reaches the expected target, and the adjustment is completed. An optical spectrum analyzer is needed to be used and is connected at a monitoring hole of an optical amplifier. The on-line upgrading capacity expansion power adjusting method of the wavelength division multiplexing optical transmission system can accurately adjust the transmitting optical power of a newly added high-speed wave channel, is implemented through the monitoring hole of the optical amplifier, and does not interrupt the original wave channel service, thus being simple and practical, solving the key problem of upgrading capacity expansion of a low-speed WDM system, and laying a good foundation for the application of the WDM system in the upgrading capacity expansion.

The invention discloses a front edge design method for optimizing a suction surface wave system and suppressing fan shock wave noise. The front edge design method comprises an element blade front edgeoptimum design method, a three-dimensional fan / rotor blade design method based on element blade front edge optimum design and a three-dimensional fan / rotor blade design method through combination ofelement blade front edge optimum design and bending modeling. On the existing ultrasonic blade basis, the shape of the front edge and the shape of a suction surface are fast changed, air flow turningangle distribution of the suction surface between the front edge point and the point E is reasonably distributed, and therefore distribution of compression waves and expansion waves on the suction surface is optimized, the interference function of the wave system and forward-stretching shock waves on the suction surface is improved, and then the effect of reducing the transonic fan shock wave noise is achieved; and the shock wave noise of an ultrasonic blade of an original round front edge can be effectively reduced by about 1.5-3.5 dB, the method can be combined with the bending three-dimensional modeling technology, the noise reduction effects of a near field and a far field are achieved at the same time, and high engineering application value is achieved.

The invention discloses a raft wave energy generating set with a flexible structure. The raft wave energy generating set with the flexible structure comprises a turbine generator, an anchor rod, and an open type cavity made of a flexible material, wherein a bellmouth-shaped water inlet section is arranged at one end of the open type cavity, the turbine generator is arranged at the other end of the open type cavity, and the open type cavity is composed of a raft impulse forming region, an expansion wave energy collecting region and a wave energy converting region which are in sequential through connection from the water inlet section to a motor mounting region; the cross section of the expansion wave energy collecting region transits from large to small to the wave energy converting region from the raft impulse forming region; the motor mounting region is positioned at the tail end of the wave energy converting region; and the flowing of flow water drives the blades of the turbine generator to rotate, and further is converted into electric energy. The raft wave energy generating set provided by the invention has a simple structure, energy converting links are less, wave energy use ratio is improved greatly, the raft wave energy generating set also can be arranged in the front of a mole, so as to generate power on one aspect, and play a wave eliminating effect on the other aspect, and the wave-block pressure of the mole is alleviated greatly.

The invention relates to an air supersonic-velocity condensation and cyclone separation spray pipe. The air supersonic-velocity condensation and cyclone separation spray pipe mainly comprises a central body, a cyclone blade, a contraction section and an expansion section, and is mainly applied to air purification, particularly the field of natural gas dehydration and heavy hydrocarbon separation.The cyclone blade is designed by an elliptical blade method, and can effectively strength cyclone; the central body can effectively reduce eddy dissipation, and the curve of the contraction section can guarantee that the velocity of air flow can be stably increased to sound velocity; the contraction section has a gradually expanded structure, and simple linear design in the conventional device isabandoned; and the curve design can avoid the influence of expansion waves, and the spray pipe is good in aerodynamic performance. In the spray pipe, the velocity of the air flow is supersonic velocity, the residence time is very short, hydrate is avoided, and chemical agents and relative recovery systems are not needed to be added. Moreover, the spray pipe has the advantages of supporting unmanned operation, along with no movement component, compact structure, small occupied space and the like, is particularly suitable for development of oil and gas fields of boundary and desert, and offshore oil and gas fields, and has a wide application range.

An ultrahigh-speed reentry test airflow ground simulation test device comprises a gas source system, a detonation section and the like. The gas source system fills hydrogen and oxygen into the detonation section through the gas conveying pipeline, detonation waves are formed after the hydrogen and oxygen mixture is detonated, the detonation waves break through a film, a first main shock wave and acenter expansion wave are formed in the shock tube, and test gas in the shock tube is compressed and accelerated by the main shock wave. The main shock wave propagates downstream to the secondary membrane in the test gas and breaks through the secondary membrane, then a second incident shock wave and a central expansion wave are formed in an acceleration section, the test gas is continuously accelerated to move downstream through the unsteady expansion effect of the second central expansion wave, and the test gas is further accelerated in the spray pipe through the steady expansion effect. Test gas at an outlet of the spray pipe can reach a reentry speed condition through the triple acceleration effect, and a vacuum system integrally pumps the part from the acceleration section to the test and pressure relief section to a high vacuum state. According to the invention, the ground simulation requirement of a high-enthalpy flow environment during ultra-high-speed reentry flight of an aircraft is met.

The invention discloses a blade profile optimization method based on a quadratic function and capable of inhibiting shock wave noise of a fan. The method comprises the steps of two-dimensional blade profile optimization and three-dimensional blade profile optimization; the shape of a leading edge and the shape of a suction surface are optimized, the shock wave noise of the ultrasonic blade profile is reduced, meanwhile, the pneumatic performance is improved, and the thickness of the leading edge is considered to ensure the requirement of the structural strength; and the change rule of the action range and the amplitude of the quadratic function in the unfolding direction is reasonably set to adapt to the change of the blade profile thickness and the incoming flow condition so as to realize the smoothness and continuity of different blade height motif stages in a three-dimensional system. According to the method, a one-variable quadratic function is introduced into a shape function of a blade profile numerical expression so that the thickness distribution of the suction surface in the fitting range can be effectively changed, the total amount of the limit mach point forward expansion waves can be increased, the strength of the forward-extension shock waves can be weakened, and therefore the shock wave noise can be reduced, the shock wave noise of the ultrasonic blade profile or a cross-acoustic rotor can be effectively reduced by about 2-3dB, and the efficiency of the cross-acoustic rotor can be effectively improved by about 0.3%.

The invention discloses a supersonic annular wing design method, and relates to the technical fields of aeronautics and astronautics. The supersonic annular wing design method is characterized in thatbased on the design of an annular Bossman wing, an annular ReZEL wing is proposed in combination with the design of a ReZEL wing. The design of the ReZEL wing is to skive an upper triangular unit tobe thin and thicken a lower triangular unit, so that an expansion wave position of a shock wave between double wings reaches a design state; four connecting supports are located at positions which areat 45 degrees with a horizontal plane and a vertical plane of the annular wing; the four connecting supports have a same radius; and the connecting supports are used for connecting an outer annular wing and an inner annular wing. During a cruise phase, the annular ReZEL wing can generate lift force at zero angle of attack and have greater lift force at different angles of attack; the lift-to-dragratio is significantly increased, so that better flight performance can be obtained.

A method and apparatus is disclosed for producing propulsion underwater with minimal acoustical emission. In basic concept, the method comprises the expulsion and sucking of liquid into and out of a liquid port of a watercraft in a manner generating compression and expansion waves adjacent the liquid port. Such expansion and compression waves generate a positive net thrust on the watercraft in a direction opposite that of their expulsion. Also disclosed is a method and apparatus for reducing the drag caused by hull skin friction by actively increasing the momentum of the liquid stream boundary layer adjacent a watercraft via the expulsion of fluid (liquid or gas) from ports in the hull of such a watercraft.

A blade capable of weakening shock strength of a transonic speed turbine relates to a blade and aims at solving the problem that the strong shock existing in a blade cascade flow field of a transonic speed air cooling turbine causes the unstable flow of the blade cascade flow field of the turbine engine and high noise of the turbine engine. A shock refracted ray combined by shock wave refraction points on a suction surface of the blade is provided with at least one cold air jet orifice communicated with an inner cavity. The blade conducts cold air jet through the cold air jet orifice, thus changing the pressure distribution in a transonic speed turbine stage blade cascade flow passage, namely after injecting cold air with relatively high pressure, increasing the local pressure value of fluid behind throat, weakening the development of expansion wave of a stationary blade cascade expansion section and further controlling the strength of strong compressional wave induced by the expansion; and simultaneously, the cold air jet on a shock propagation passage weakens the local pressure difference and the propagation of shock wave toward the suction surface, thus leading the flow of the blade cascade flow field to be more stable and reducing the noise of the turbine engine.

Devices are provided for regulating the pressure pulses and flow pulses that are emitted in the exhausts of apparatus and equipment such as internal combustion engines, compressors and poppet valves. The pressure pulses are reflected from a reflecting member to their source as expansion waves that create a reduced pressure zone that allows subsequently emitted flow pulses to be more easily emitted, thereby enhancing the operational efficiency of the apparatus and equipment. Flow pulses that contain exhaust products such as the products of combustion from an internal combustion engine are regulated so that the exhaust flow they generate is not unduly restricted within the device and an undesirable back pressure that resists the emission of flow pulses from their source created. In regulating the pressure and flow pulses, the devices attenuate the noise generated by the pressure and flow pulses.

The invention provides a high-speed railway tunnel structure. The high-speed railway tunnel structure is characterized in that both side surfaces or the top of a tunnel lining is provided with openings which are communicated with the atmosphere, 3-5 openings are distributed in the full length direction of the tunnel, the neighboring openings are spaced at equal intervals, and the opening close to the portal of the tunnel is 20-50 m away from the portal. According to the high-speed railway tunnel structure, 3-5 the openings are arranged on the side walls or the top of the tunnel lining, so that the transient pressure inside a locomotive and the micropressure peak value of the portal can be reduced, the micropressure wave effects of the portal can be released, and the locomotive additional air resistance of a long tunnel can be reduced obviously; meanwhile, the energy of expansion waves reflected back to the tunnel from the exit of the tunnel can be reduced, further the negative pressure inside opencuts and on the outer wall of the locomotive can be effectively released, and the fatigue damage of the opencut structure and auxiliary facilities due to aerodynamic load can be reduced. Besides, the high-speed railway tunnel structure is restricted little by surrounding terrains and has high adaptability, and meanwhile, is low in construction cost and highly practical.

The invention discloses a design method of an ejector nozzle experimental device for simulating outflow of an aircraft. According to the method, an initial molded line of the meridian plane of the outflow nozzle under a specific flight Mach number is designed by adopting a characteristic line method, then an outer flow channel asymmetric nozzle meridian plane molded line meeting requirements is obtained by adjusting an outer flow channel symmetric factor G and a throat channel height h, and an outer flow channel three-dimensional model is further obtained; and finally, an outflow pressure stabilizing section and a circumferentially distributed air inlet section are designed. According to the design method, the outflow channel is designed through the characteristic line method, original experiment assemblies can be utilized to a large extent, the method is simple, programs are easy to achieve, and the profile of the outflow channel under different mach numbers can be designed. The supersonic airflow generated by adopting the design method does not have expansion waves or shock waves, the flow parameters are uniform, the outflow environment of the aircraft can be accurately simulatedon the premise of ensuring the economical efficiency, and a feasible experimental device design method is provided for researching the inflow and outflow coupling mechanism of the high-altitude flight working condition of the aircraft.

An optical waveguide structure (301) comprises at least one first substrate (11) and an expansion wave guidance (20), the first substrate (11) is connected with the expansion wave guidance (20), the expansion wave guidance (20) comprises a plurality of semipermeable membranes (a, b, c, d, e) mutually obliquely laid into the substrate (22) in parallel, the at least one side of the left and right sides of the optical waveguide (301) is provided with light absorption materials, wherein incident light (L1, L2) enter the expansion wave guidance (20) and then emits a plurality of emitting light (L2,L3, L4, L5, L6, T2, T3, T4, T5, T6) being mutually in parallel; when the incident light (L1, L2) enters the expansion wave guidance (20) and is reflected by at least one of the semipermeable membranes (a, b, c, d, e), the reflection light enters the first substrate (11), and then is subjected to total reflection spreading through the first substrate (11) to the side, provided with the light absorption materials, of the optical wave guidance (301), or the incident light (L1, L2) enters the expansion wave guidance (20) and is reflected by at least one of the semipermeable membranes (a, b, c, d,e), the reflection light spreads to the side provided with the light absorption materials of the optical wave guidance (301) through the first substrate (11). The optical system comprising the optical wave guidance (301) can reduce the ghost image so as to improve the imaging quality.

The invention discloses a method for rapidly determining curved shock wave compression flow field air flow parameters. According to the method, change of flowing parameters caused by a compression wave system emitted by a compression wall surface and change of flowing parameters caused by an expansion wave system reflected by curved shock waves are calculated independently, in the calculation process, the compression wave system in an original flow field is simplified into Prandtl-Meyer compression waves gathered on the point S, the expansion wave system in the original flow field is simplified into an expansion wave passing through the point S, the compression wall surface parameter, the curved shock wave shape, parameters after waves, the shape of streamlines in the flow field, parameters on the streamlines and an output parameter are determined rapidly according to the simplified Prandtl-Meyer compression waves and the expansion wave. According to the method for rapidly determining curved shock wave compression flow field air flow parameters, the calculation process is simple, the calculation result can reach high precision, and the method can be used for calculating and analyzing the flow field.

The invention discloses a supersonic turbine nozzle blade and a design method thereof. The blade profile of each nozzle blade is composed of two curves by taking a leading edge and a tail edge as division points, one of the curves of each nozzle blade is a pressure surface curve, the other one of the curves of each nozzle blade is a suction surface curve, each pressure surface curve is provided with a corresponding convex point which is a corresponding first convex point, each suction surface curve is provided with a corresponding convex point which is a corresponding second convex point, andthe first convex point on each nozzle blade and the second convex point on the adjacent nozzle blade are aligned and are located at the same position perpendicular to a streamline to form a throat part of a nozzle runner, so that a corresponding convergence channel is formed by the two adjacent nozzle blades at the upstream of the throat part, the two adjacent nozzle blades form a corresponding symmetrical expansion channel at the downstream of the throat part, and accordingly, airflow is accelerated from subsonic flow at the airflow inlet side to sonic flow and then accelerated to supersonicflow. According to the nozzle blade, a supersonic flow expansion wave system can be distributed more gently, supersonic flow loss is reduced, and meanwhile the expansion capacity of a nozzle is improved.