18 results about "Expansion wave" patented technology

A system for trapping a target condensable gas may include a supersonic condensation nozzle configured to receive a gas stream containing a target condensable gas component and accelerate the gas into a supersonic gas stream; at least one particle injector in the supersonic nozzle for injecting particles such that the particles are in the supersonic gas stream, whereby a target gas condensate is formed on a portion of the particles; an expansion device defining a channel having a lobe, causing the flow of particles containing the target gas condensate to generate an expansion wave to separate the flow into a first flow path in which a concentration of particles containing the target gas condensate increases and a second flow path in which a target condensable gas content of the gas decreases; and a diverter having at least a first channel, the diverter configured to collect the first flow path.

The application relates to the field of pipe materials, in particular to a PE and steel skeleton composite pipe and a pipe material connecting structure. The pipe body is provided with a socket end and a spigot end fixedly arranged at two ends of the pipe body respectively, the socket end is used for supporting the spigot end of another pipe material, at least one expansion wave is arranged in the socket end, an inner groove of the expansion wave is used for mounting a sealing ring, the outer wall of the spigot section can be matched with the inner ring of the sealing ring, a first pair of locking pieces are fixedly arranged on the socket end, a second pair of locking pieces capable of being fixedly locked with the first pair of locking pieces of another pipe material are fixedly arranged on the pipe body, the spigot end can be axially pressed against the socket end or the pipe body of another pipe material, the axial position of the first pair of locking pieces relative to the second pair of locking pieces of another pipe material is relatively adjustably arranged, so that the stretching degree of the expansion wave can be adjusted to adjust the tightness of the sealing ring. The application provides the PE and steel skeleton composite pipe and the pipe material connecting structure, the pipe material and the connecting structure are convenient to install, have good sealing performance and the tightness of the sealing ring can be adjusted.

The invention relates to a combustion chamber with a wide-range ignition starting structure, an engine and a starting method. The combustion chamber comprises a combustion chamber body and the ignition starting structure arranged on the combustion chamber body. A flame stabilizing concave cavity is formed in the first side wall of the combustion chamber body, and the ignition starting structure is arranged on the second side wall opposite to the first side wall. The ignition starting structure comprises at least one shock wave generator; the shock wave generator comprises a generator cylinder with a fan-shaped section and a driver; and the driver is used for driving the generator cylinder to deflect so as to adjust the relative angle between the side plane of the generator cylinder and the supersonic incoming flow, and is used for inducing generation of oblique shock waves, expansion waves and reattachment shock waves in the supersonic incoming flow. According to the scheme, the ignition starting problem of the engine under the conditions of combustion chamber incoming flow and wide-range engine oil temperature change is effectively solved.

The invention provides a compressible flow field structure identification method based on local convection characteristic distribution, and the method comprises the steps: calculating a nonlinear convection characteristic value vector along three space orthogonal squares according to a flow field variable on each point or unit for an overall calculation region; for any point or unit, according to the flow field variables around the point or unit, calculating a gradient matrix of the convection characteristic value in the space orthogonal direction; according to the gradient matrix, calculating an index value of the vortex, an index value of the shock wave, an index value of the expansion wave and an index value of the separation structure; and outputting the flow field junction to which each point or unit belongs according to a compressible flow field classification criterion. The method provided by the invention has parameter freedom and good anti-noise characteristic, and has an identification and capture effect superior to that of a traditional method for a noise-containing or high-dissipation flow structure in simulation / experiment data.

This invention provides a method for predicting the waveform of a wave passing through a branch tunnel, a waveform prediction device, and a waveform prediction program for a wave passing through a branch tunnel, which can predict the waveform of the wave passing through the branch tunnel quickly and accurately. [Solution] When a train 1 moving through the main tunnel 3A passes through the branch section 3C where the branch tunnel 3B branches off from the main tunnel 3A, a branch tunnel passage wave W is generated. 31 ,W 32 A method for predicting the waveform of a branch tunnel-passing wave, wherein the waveform of the branch tunnel-passing wave W is predicted by a compact Green's function based on the cross-sectional area change rate distribution of the front of train 1. 31 ,W 32 The process includes a waveform prediction step to predict the waveform of the compressed wave passing through the main tunnel 3A and the waveform of the wave passing through the branch tunnel W. 31 The process includes predicting the waveform of the expansion wave passing through branch shaft 3B as the branch shaft passing wave W. 32 This process includes predicting the waveform.

An improved hydrogen generation system comprising a multi-port wave reformer in which shock and expansion waves are created in a manner causing head-on colliding shock waves and multi-stage compression where reacting gases within a eight port wave reformer are heated and compressed to thermally crack or decompose one or more fuel sources, such as hydrocarbon fuels, to generate a fuel product containing hydrogen.

The invention discloses an inward rotation type air inlet channel overflow port which comprises a rounding section and two straight front edge sections, the two straight front edge sections are smoothly and tangentially connected with the two ends of the rounding section respectively, two smooth protrusions are locally arranged on the rounding section, one protrusion is located on the downstream of the tangent point of the rounding section and the straight front edge sections, and the other protrusion is located on the downstream of the tangent point of the rounding section and the straight front edge sections. The other protrusion is located on the downstream of the tangent point of the rounding section and the other straight leading edge section, the highest points of the protrusions are located on the upstream of the maximum heat flow peak position of the wall face of the rounding section, then the wall face between the tangent point and the highest points of the protrusions is lifted, and therefore compression waves intersect with detached shock waves of the straight leading edge section on one side in advance. Expansion waves entering and exiting from the wall face of the rounding section are matched earlier, the maximum heat flow peak value of the downstream of the highest point of the protrusion can be further weakened, and the heat flow peak value of the wall face of the rounding section can be reduced.

The invention discloses a combustion chamber wall surface structure for inhibiting oblique detonation wave induced boundary layer separation, which comprises a plurality of columns of induction units arranged on an oblique wall at a top inflection point of an engine combustion chamber, and the plurality of columns of induction units are arranged on an oblique surface of the oblique wall along the length direction of the oblique wall; each row of induction units comprises a plurality of scale reflecting plates which are arranged side by side in the width direction of the inclined wall, each scale reflecting plate is in a table shape and is provided with a lower bottom surface, an upper bottom surface and an inclined side surface, the included angle between each inclined side surface and the lower bottom surface is an obtuse angle, and the inclined side surface on one side of each scale reflecting plate is fixed to the inclined wall; the multiple columns of induction units are divided into odd columns of induction units and even columns of induction units, and the scale reflecting plates in the odd columns of induction units and the scale reflecting plates in the even columns of induction units are arranged in a staggered mode; a plurality of rows of induction units are additionally arranged on the smooth wall face of the combustion chamber to form a fish scale type induction structure, and expansion waves formed by the fish scale type induction structure are used for continuously weakening and restraining boundary layer separation.

The invention provides a variant hypersonic gliding aircraft based on high-pressure capture wings. The variant hypersonic gliding aircraft comprises a revolution body, the high-pressure capture wings and a connecting device. A reference configuration of the aircraft, namely a lower boundary of a front profile of a variant based on a flight Mach number, is taken as a design point condition, the design point condition is designed according to a design principle of a high-pressure capture wing, and a revolution body is generated through rotation by taking a von Karman curve as a generatrix; the high-pressure capturing wing is mounted at the upper part of the rear end of the rotating body machine body through a connecting device; body shock waves and expansion waves of the revolution body are in axial symmetry, and the high-pressure capturing wings are of a flat plate structure integrally in a dovetail shape. The connecting device is of a vertical plate structure with the horizontal position and the height position adjustable. The aircraft is simple and reasonable in structural design, the capturing wings are in variant design, the positions of the capturing wings are reasonably adjusted according to different incoming flow Mach number conditions, the high-pressure capturing capacity of the capturing wings can be utilized to the maximum extent, and the aerodynamic performance of the aircraft under the wide Mach number condition is effectively improved.

The system for capturing target condensed gas includes a supersonic condensation nozzle configured to receive a gas flow having a predetermined target condensed gas content and for accelerating the gas into a supersonic gas flow; at least one particle injection device within the supersonic nozzle for injecting particles into the supersonic gas flow, whereby target gas condensate is formed on some of the particles; an expansion device that generates an expansion wave in the flow of particles containing the target gas condensate, defining a conduit with a convex corner that separates the flow into a first flow path having particles with a high concentration of target gas condensate and a second flow path of gas with a lower target condensed gas content; and a flow divider having at least the first conduit, the flow divider configured to receive the first flow path.

The invention discloses a hyperbolic cone configuration shock wave interference type conversion heat reduction design method based on back cone curve regulation and control, relates to the technical field of aerodynamic heat protection of hypersonic aircrafts, and aims to solve the problem of extreme thermal load of a wall surface of a hyperbolic cone configuration caused by fourth type (Type IV) shock wave interference. Giving a hyperbolic cone wall surface curve, a flight Mach number and incoming flow conditions, and solving shock wave parameters through a bending shock wave theory; fitting curvilinear equations of the two shock waves, solving intersection points by combining the curvilinear equations of the shock waves, and determining positions of three wave points; based on a Rankine-Hugoniot relational expression and an Eddy shock wave interference classification criterion, establishing a shock wave interference type discrimination model; a rear cone curve is changed to realize the conversion from strong interference containing ultrasonic rapid jet flow to weak interference containing weak expansion wave, so that the shock wave interference type is directionally converted from the fourth type to the sixth type, and the wall surface heat flow is obviously reduced.

The utility model discloses an airborne lighting bomb which comprises an empennage component (1), a fuse (2), a lighting component (3) and a hood component (4). The front end of the empennage part (1) is in threaded connection with the rear end of the lighting part (3), the front end of the lighting part (3) is in threaded connection with the rear end of the hood part (4), and the fuse is arranged in an inner cavity of the empennage part (1) and used for igniting the lighting part (3); when the lighting bomb is not used, the empennage part (1) is fixed to be in a folded state, and after the lighting bomb is air-dropped, the empennage part (1) is unfolded. By reducing the size of the tail part, reducing the shrinkage ratio of the tail part, increasing the slenderness ratio of the bomb body and the tail part and weakening airflow separation and expansion wave intensity, the purpose of reducing the resistance of the tail part is achieved, the rotating speed in the lighting distance combustion process is conveniently reduced, conditions are created for normal combustion of the lighting distance, and the lighting distance combustion device has the advantages of being simple in structure, small, light, high in operability and the like.

The present application relates to the technical field of aviation, and in particular to a supersonic civil aircraft configuration. In the supersonic civil aircraft configuration of the present application, engine nacelles are mounted above a rear fuselage between wings and horizontal stabilizers; the lower portions of inlets are shielded by the trailing edges of the wings, and the lower portions of outlets are shielded by the leading edges of the horizontal stabilizers; by means of the shielding effect of the wings and the horizontal stabilizers, shock waves generated by engines are prevented from propagating to the ground, effectively reducing the ground sonic boom loudness of such an aircraft; the engines and pylons can perform small-angle pitch deflection around pivot shafts in the pylons, thereby adjusting mounting angles of the engines and improving the propulsive efficiency of the engines under off-design operating conditions; during supersonic cruising, the tips of the horizontal stabilizers can be folded upward to provide an additional vertical stabilizer volume, so as to compensate for the loss of vertical stabilizer efficiency during supersonic cruising, and preventing expansion waves generated by engine exhaust nozzles and the horizontal stabilizers from propagating to the ground; and during low-speed takeoff / landing, the tips are unfolded to ensure a large vertical stabilizer area, so as to meet the requirements of trim in takeoff / landing configuration.

The application belongs to the field of aircraft design and relates to a low-wave resistance aerodynamic configuration and design method of an aircraft, which comprises a body and a wing, and the body and the wing are both rotary bodies, the body comprises a front cone and a rear cone; incoming airflow is compressed for the first time at a first compression cavity on the front cone, the incoming airflow is compressed for the second time after reaching a second compression cavity, a front shock wave surface is formed, and then the incoming airflow forms a body expansion wave surface when reaching a decompression cavity at the rear cone, and the decompression is realized under the action of the shock wave expansion wave; in this way, the pressure acting on the front cone of the body only undergoes one compression, and the pressure acting on the rear cone of the body undergoes two compressions and one decompression; therefore, the pressure acting on the rear cone of the body is close to the pressure acting on the front cone of the body, that is, the pressure difference between the front cone and the rear cone of the body is relatively small, the pressure difference between the front cone and the rear cone of the body can be significantly reduced, and thus the shock wave resistance of the aerodynamic configuration is greatly reduced.

The present application relates to a combustion chamber with wide range ignition starting structure, an engine and a starting method, wherein the combustion chamber comprises: a combustion chamber body, an ignition starting structure arranged on the combustion chamber body; a first side wall of the combustion chamber body is provided with a flame stabilizing cavity, and the ignition starting structure is arranged on a second side wall opposite to the first side wall; the ignition starting structure comprises: at least one shock wave generator; the shock wave generator comprises: a fan-shaped generator column body, and a driver; the driver is used for driving the generator column body to deflect, so as to adjust the relative angle between the side plane of the generator column body and the supersonic speed airflow, and is used for inducing the generation of oblique shock wave, expansion wave and reattachment shock wave in the supersonic speed airflow. The scheme effectively solves the ignition starting problem of the engine under the condition that the combustion chamber airflow and the engine oil temperature change in a wide range.

The application belongs to the field of aviation technology, and particularly relates to a supersonic civil aircraft layout. The supersonic civil aircraft layout of the application is characterized in that an engine nacelle is installed above the rear fuselage between a wing and a horizontal tail, an air inlet is shielded by a wing trailing edge below, an air outlet is shielded by a horizontal tail leading edge below, shock waves generated by the engine are prevented from spreading to the ground by the shielding effect of the wing and the horizontal tail, and the ground sound blast loudness of the aircraft is effectively reduced; the engine and the pylon can be deflected by a small angle around a rotating shaft in the pylon to adjust the installation angle of the engine and improve the propulsion efficiency of the engine under non-design point working conditions; the wing tip of the horizontal tail can be folded upward to provide additional vertical tail capacity when the aircraft cruises at supersonic speed, compensate for the loss of vertical tail efficiency when the aircraft cruises at supersonic speed, and prevent the spread of expansion waves generated by the engine tail nozzle and the horizontal tail to the ground; the wing tip is lowered when the aircraft takes off and lands at low speed to ensure a large horizontal tail area to meet the trim requirements of the take-off and landing configuration.

This invention relates to a high-efficiency cooling system for supersonic turbine blade tips based on shock wave interference effects. Specifically, it involves setting film cooling holes of different sizes in the leading and trailing edge regions of the turbine blade. Through numerical simulation, the wave system structure of the blade tip clearance is accurately captured, clarifying the distribution of high-pressure / low-pressure regions formed by shock wave interference. The relationship between shock waves and tip heat transfer and flow field is precisely analyzed, as well as the flow field distribution within the clearance, which differs from conventional subsonic blade tip clearances, due to the complex wave system structure of shock waves and expansion waves. By arranging film cooling holes of different diameters and spacings close to the pressure side in the leading and trailing edge regions, the non-uniform spatial layout of the cooling holes is optimized, greatly improving cooling efficiency and targeting. This solves the cooling failure problem caused by neglecting the influence of shock wave interference, achieving precise matching between flow field characteristics and cooling structure design, resulting in greater targeting.

The application discloses a combustion chamber wall surface structure for inhibiting oblique detonation wave-induced boundary layer separation, comprising multiple columns of induction units, which are arranged on an oblique wall at a top corner of an engine combustion chamber, and the multiple columns of induction units are arranged on the slope of the oblique wall along the length direction of the oblique wall, each column of induction units comprises multiple scale reflection plates arranged side by side along the width direction of the oblique wall, the scale reflection plate is in the shape of a platform body, has a lower bottom surface, an upper bottom surface and an inclined side surface, and the included angle between the inclined side surface and the lower bottom surface is an obtuse angle, one side of each scale reflection plate is fixed with the oblique wall, the multiple columns of induction units are divided into odd-numbered column induction units and even-numbered column induction units, the scale reflection plates in the odd-numbered column induction units and the scale reflection plates in the even-numbered column induction units are arranged alternately, and the smooth wall surface of the combustion chamber is additionally provided with the multiple columns of induction units to form a scale type induction structure, and the expansion wave formed by the scale type induction structure continuously weakens and inhibits the boundary layer separation.