13 results about "Sonic boom" patented technology

The invention relates to the technical field of aerodynamic configuration design of aircrafts, in particular to an aircraft low sound detonation reverse design method based on a transfer learning multi-fidelity neural network. According to the technical scheme, the near-field signal of the initial scheme is modified, the target near-field signal with smaller sound explosion is obtained through iteration in combination with the optimization method, and then the aircraft profile meeting the target near-field signal is obtained through the inverse design method. According to the method, the solvability of a generalized Burgers equation is adopted, a proxy optimization method is applied according to an initial scheme, and low-sound-explosion target near-field overvoltage distribution is determined; a multi-fidelity deep neural network is constructed based on transfer learning, the characteristic of low demand of fine tuning on data volume is utilized, the mapping relation between the aircraft profile and near-field overvoltage distribution can still be accurately captured when high-fidelity data is less, and due to the fact that the transfer learning does not need to be connected in series with multiple models, the accuracy of the multi-fidelity deep neural network is improved. Error accumulation caused by insufficient training of a single model during traditional multi-fidelity modeling can be avoided to a certain extent.

The embodiment of the invention relates to a sound output device and an audio data interruption protection method of the sound output device, which are used for driving loudspeaker equipment. The sound output device comprises a first-in first-out buffer used for receiving and outputting audio data; the buffer circuit is used for temporarily storing the received audio data; and an explosion-proof sound protection device. During normal operation, the explosion-proof sound protection device receives audio data from the buffer circuit and drives loudspeaker equipment. When an underflow notification of the first-in first-out buffer is received, the first-in first-out buffer acquires the last audio data from the buffer circuit and gradually reduces the volume of the audio data so as to avoid the generation of sonic boom. The design can effectively prevent sudden sonic boom caused by audio data interruption, and protect loudspeaking equipment and hearing of a user.

The invention provides a large-scale supersonic civil aircraft wing body fusion low-sonic-boom aerodynamic layout, and relates to the field of supersonic civil aircraft design, the large-scale supersonic civil aircraft wing body fusion low-sonic-boom aerodynamic layout comprises a nose, a middle fuselage, a wing body transition section, wings and a V-shaped empennage; the machine head is in a duckbilled flat shape, the upper contour line is a downward-inclined curve, the lower contour line is an upward-warping curve, the side contour line is a sharp-arch-shaped curve, the section of the head is in the shape of a combination of a triangle and an arc, the upper portion of the section of the head is a triangle with the top in arc transition, and the lower portion of the section of the head is an arc. The camber of the center line of the middle fuselage is small, and only the front section bends down slightly; the middle fuselage is connected with the roots of the wings through wing body transition sections; the wings are arranged on the two sides of the middle fuselage and adopt an arrow-shaped wing layout with a small aspect ratio, a variable dihedral angle and a variable torsion angle and a camber wing type; the V-shaped empennages are arranged on two sides of the tail end of the According to the method, the problem that a gap still exists in the fusion low-sonic-boom aerodynamic layout of the wing body of the large supersonic speed civilian aircraft is solved, and reference is provided for research on the fusion layout of the supersonic speed wing body.

Commercial supersonic aircraft and associated system and methods. A representative real time sonic boom warning system can detect when the trajectory of the commercial supersonic aircraft is likely to cause sonic booms that disturb designated areas (e.g., over land or population centers) and notify the operator accordingly.

The invention provides a test system and method for simulating turbulence and sonic boom interference based on a small shock tube, and belongs to the technical field of sonic boom ground simulation tests. The objective of the invention is to solve the problem that a simulation result is deviated from reality due to the fact that an acoustic explosion model does not simulate an atmospheric turbulence environment in the prior art. A high-voltage system, a high-voltage driving section, a double-film section, a low-voltage driven section and a vacuum system are connected in sequence, and a speed measurement piezoelectric sensor is installed on the side wall of the low-voltage driven section; the grid turbulence disturber is arranged on the side wall of the low-pressure driven section; the turbulence control system is connected with the grid turbulence disturber; the hot-wire probe is arranged at the front end of the hot-wire anemometer and is arranged on the side wall of the low-voltage driven section; the test board is provided with a PSP coating, and the test board, the schlieren high-speed camera and the PSP high-speed camera are all mounted on the side wall of the low-voltage driven section; the test board is provided with a high-frequency pressure sensor. The problem that in the prior art, a sound explosion model does not simulate the atmospheric turbulence environment, and consequently the simulation result deviates from the reality is solved.

The application provides a flow field overpressure distribution nonlinear reduction prediction method, which is suitable for nonlinear dimension reduction and rapid prediction of flow field overpressure distribution data, and aims to overcome the limitation that traditional linear dimension reduction method is difficult to process high-dimensional complex aerodynamic data. The method firstly generates a design variable sample set based on the parameterized shape of the aircraft, and obtains corresponding high-dimensional overpressure distribution data through CFD simulation; then the overpressure distribution data is nonlinearly reduced by using a Gaussian process hidden variable model to extract hidden space variables and reconstruct the model; then the mapping relationship between the design variables and the hidden space variables is established by using a Kriging surrogate model, so that the rapid prediction from the design variables to the overpressure distribution is realized. The method does not need to solve the flow control equation, and can complete the reconstruction and prediction of the overpressure distribution within seconds, significantly improving the efficiency of the aircraft aerodynamic layout design, and is suitable for rapid design and evaluation of high-performance aircraft such as low sonic boom aircraft.

This application relates to a method, apparatus, device, and medium for locating sonic booms of a single rocket debris. The method includes: calculating the time when the monitoring equipment receives the sonic boom signal based on the distance between the monitoring equipment and the sonic boom source and the time when the rocket debris generates the sonic boom signal; calculating the squared difference between the sonic boom signal received by each monitoring equipment and the predicted sonic boom time, setting the squared difference as the objective function of the sonic boom location model, setting constraints on the sonic boom location model based on the range of the sonic boom's occurrence time and location, constructing the sonic boom location model using the objective function and constraints; optimizing the hyperparameters of a genetic algorithm using a Bayesian optimization method, and solving the sonic boom location model using the optimized genetic algorithm to obtain the sonic boom location result. This method can accurately locate rocket debris.

The invention relates to the field of hearing aids, in particular to an impulse noise suppression method and system for a hearing aid, and the system comprises a machine body module, a processing module, a protection module and a driving module. The hearing aid main body comprises a shell, a sound receiver, and a battery, a charging interface, a processor, a broadcaster and keys which are arranged in the sound receiver, the external ear hanging piece is hung at the external ear part and is used for mounting and supporting the hearing aid main body, and sudden pulse noise is inhibited by setting the maximum upper limit of the volume when the player plays; the driving module drives the sound receiving device to rotate, so that the follow-up sound processing module can process pulse noise conveniently, meanwhile, the driving module can drive the protective net to wrap the sound receiving device, at the moment, the protective net can reduce sonic boom caused by air flow impact when the sound receiving device receives sound and filter some unnecessary noise to a certain extent, external pulse noise interference is further prevented, and the sound receiving device can be used for sound receiving. Therefore, the hearing aid can be normally used under the interference of pulse noise.

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 utility model relates to the technical field of industrial dust removal equipment, and discloses efficient dust suppression equipment for a drag chain distribution vehicle, when the distribution vehicle moves to a blanking area along a guide rail, a side discharge device transports materials on the distribution vehicle to a coke blanking box in the blanking area, then double-fluid drag chain equipment is started, and the materials are discharged from the coke blanking box. The double-fluid spray head enables gas in the gas-liquid double pipe to form a local supersonic speed area at the throat part of the spray nozzle, at the moment, strong turbulence and shearing force can be caused to form a sonic boom effect, and then water molecules in the gas-liquid double pipe are torn into tiny liquid drops (generally smaller than 10 microns), so that dust in the transportation process is inhibited, and the transportation efficiency is improved. And no extra dust pollution is generated.

1. Name of the designed product: game machine (sonic boom). 2. Use of the designed product: for entertainment equipment. 3. Design points of the designed product: in shape. 4. Picture or photo that best shows the design points: perspective view.

A non-invasive shock wave transdermal injection device for cosmetic use relates to the field of cosmetic technology and includes a housing, a charging mechanism, a pilot solenoid valve, a shock wave generator, and a shock wave balancing tube. One end of the pilot solenoid valve is connected to an air source, and the other end is connected to the shock wave generator. The shock wave generator is provided with a front end of a shock wave generating chamber that gradually expands along the direction of the airflow. The shock wave balancing tube is connected to the shock wave generator and is provided with a rear end of the shock wave generating chamber that gradually contracts along the direction of the airflow and an injection chamber connected to the rear end of the shock wave generating chamber. The inner diameter of the injection chamber gradually expands along the direction of the airflow. The pilot solenoid valve is used to accelerate gas and sequentially inject it through the front end of the shock wave generating chamber, the rear end of the shock wave generating chamber, and the injection chamber. The shock wave balancing tube is connected to the charging mechanism. The shock wave transdermal injection device can generate shock waves, and the shock waves reflect and focus the energy of the sound wave to form a sonic boom. The shock wave gas jet effect truly achieves the effect of penetrating the skin, which is far superior to traditional injection devices.

The application belongs to the technical field of aircraft sonic boom wind tunnel test, and particularly relates to an aircraft sonic boom experimental model, a sonic boom wind tunnel test system and a method thereof, wherein the aircraft sonic boom experimental model comprises a front fuselage model, a middle fuselage model with wings, a tail fuselage model with tail wings, a tail fuselage model without tail wings, a support rod, a blocking block and a horizontal measurement platform; the rear end of the front fuselage model is detachably connected with the front end of the middle fuselage model with wings; the front end of the tail fuselage model with tail wings or the tail fuselage model without tail wings is detachably connected with the rear end of the middle fuselage model with wings; the front end of the support rod is detachably connected with the back of the middle fuselage model with wings, and the rear end is connected with a support structure; the rear end of the front fuselage model is provided with a mounting hole, and the blocking block or the horizontal measurement platform is mounted in the mounting hole.