26 results about "Sonic boom" patented technology

The invention discloses a test system and method based on electric spark simulation acoustic detonation waves, and belongs to the technical field of acoustic detonation wave ground simulation tests. The problem that in the prior art, an acoustic explosion model does not simulate the atmospheric turbulence environment is solved. The turbulence generation system is adjusted to add a disturbance source for a flow field to generate plane jet flow, the hot line measurement system is used for measuring turbulence intensity, the electric spark acoustic explosion source system is used for generating N-type wave signals used for representing typical acoustic explosion characteristics, and the acoustic explosion wave measurement system is used for measuring acoustic explosion wave signals and storing measurement data. According to the ground test simulation system and method, the N-type wave of the far-field acoustic detonation, the atmospheric turbulence and the influence between the N-type wave and the atmospheric turbulence can be completed on the ground, the influence of the atmospheric turbulence on the acoustic detonation can be considered, a ground simulation platform is provided for research on the interaction between the acoustic detonation wave and the atmospheric turbulence, and the prediction precision of the far-field acoustic detonation is improved.

The invention discloses an audio power amplifier starting-up sonic boom suppression method based on a sound masking effect, and belongs to the technical field of power amplifiers, and the method comprises the following steps: S1, presetting a voltage reference value for audio detection; s2, correspondingly outputting a suppression signal or a working signal according to the comparison between the input voltage value and the voltage reference value; s3, generating a PWM signal, and connecting the PWM signal to a switch module through an AND gate and a buffer; s4, when the suppression signal is generated, it is judged that the audio input signal has sonic boom at the moment, and the output end of the power amplifier does not generate an impact signal; and when the working signal is generated, the output end of the power amplifier generates an impact signal containing a suppression effect. According to the invention, through input signal threshold detection, the composite sound superposed with the preset switch music signal is generated, and the method has the advantages of solving the bottleneck of response lag during shock-wave suppression of the D-type power amplifier and promoting the D-type power amplifier to be miniaturized.

The invention relates to the technical field of computer processing, and discloses a sound quality detection method and device, computer equipment and a storage medium, and the method comprises the steps: obtaining to-be-detected audio data; feature detection is carried out on the to-be-detected audio data to obtain audio features, the audio features comprise cut-off frequency and / or sonic boom features, the cut-off frequency is obtained based on frequency domain data corresponding to the to-be-detected audio data, and the sonic boom features are obtained based on time domain data corresponding to the to-be-detected audio data; and determining a tone quality detection result of the to-be-detected audio data based on a size relationship between the audio features and corresponding preset thresholds.

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.

An apparatus comprising: a first power stage (PS1) having a first power stage input (output of 308) and a first power stage output (voltage VN at the PS1 output); a second power stage having a second power stage input and a second power stage output; and a first driver circuit (408) having a first driver input (having audio signals VINM and VINP) and a first driver output (output of 408), the first driver output coupled to the first power stage input (output of 308). The control circuit (310) is configured to set the first power stage output (VN) in response to the first driver input being in an inactive state (VINM and VINP being inactive). The apparatus also includes a second driver circuit having a second driver input and a second driver output, the second driver output coupled to the second power stage input, and the second driver input coupled to the first power stage output, and the second driver circuit is configured to set the second power stage output to track the first power stage output.

A method (100) is provided for detecting a firearm (10) equipped with a laser-emitting emitter; the method (100) comprises a laser detector (2); three microphones (3) configured to detect a sound; a clock; a signaler (4) configured to indicate the firing position of said firearm (10); and a board determining the firing position of the firearm (10) based on the laser of the emitter detected by the detector (2) and the detection and timing of the sonic boom and muzzle blast detected by each of the microphones (3).

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.

An embodiment of the present invention discloses a method and device for analyzing sonic boom test data, relating to the technical field of wind tunnel sonic boom tests in the field of fluid mechanics, which can improve the accuracy and efficiency of the entire sonic boom test data analysis system, reduce the number of tests, and save test resources. The present invention includes: according to the data acquisition request sent by the client, reading the measurement data of the sonic boom test through a pressure measurement rail; analyzing and processing the initial data through a probability model stored in the server; solving the model parameters by using the maximum likelihood method; obtaining the model sonic boom signal to be measured through the parameter results, and finally returning the obtained results to the client. The present invention is applicable to equipment design involving wind tunnel sonic boom tests.

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 relates to the technical field of aircraft design and acoustics, in particular to a sound detonation signal intelligent inversion method based on a conditional generative adversarial network, which comprises the following steps: S1, selecting a sound detonation reference near-field signal, and calculating a sound detonation reference far-field signal; s2, designing a sonic boom target far-field signal based on the obtained sonic boom reference far-field signal; s3, adding disturbance on the basis of the sound detonation reference near-field signal selected in the step S1 to obtain a sound detonation near-field signal after the disturbance is superimposed, calculating a corresponding sound detonation far-field signal by using the sound detonation near-field signal after the disturbance is superimposed, and establishing a sound detonation near-far-field signal sample set; s4, training a conditional generative adversarial network cGAN by using the obtained acoustic detonation near-far field signal sample set; and S5, inputting the acoustic explosion target far-field signal designed in the step S2 into a trained conditional generative adversarial network cGAN, and inverting and outputting an acoustic explosion target near-field signal. The method can get rid of the constraint of the sound detonation inversion distance, and improves the application range of sound detonation inversion.

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 discloses a method for relieving the pneumatic effect of a tunnel. The method comprises the following steps that (1) a first draught fan is arranged in the tunnel; (2) arranging a buffer chamber in the tunnel; (3) roughening the inner surface of the tunnel and the upper surface of a ballast bed in the tunnel; and (4) arranging a sound absorption structure in the tunnel. The first draught fan is arranged in the tunnel, the first draught fan can disturb airflow in the tunnel, and therefore micro-air-pressure waves generated when a train passes through the tunnel are weakened; the buffer chamber is arranged in the tunnel, the head wave gradient of pressure waves can be slowed down, the pressure waves caused when a train enters the tunnel are disturbed, then micro-air-pressure waves and sonic boom are relieved, and the pressure relief and buffer effects are achieved; the inner surface of the tunnel and the upper surface of the ballast bed are roughened, and micro-pressure waves in the tunnel can be relieved and interfered through the design; the sound absorption structure can absorb sound, reduce noise and relieve micro air pressure waves and sonic boom.

The invention discloses a method for measuring a sonic boom signal in a flight test, which comprises the following steps: obtaining a measurement position of a ground sonic boom signal according to parameters of a flight Mach number, a flight altitude and an attitude angle, reasonably selecting the position and number of measurement points, and the measuring equipment comprising a capacitive low-frequency microphone, a preamplifier, a data collector and a portable computer. At each measurement point, a plurality of capacitive low-frequency microphones are used and respectively configured with preamplifiers to form a microphone array, the microphone array electrical signal is connected to the data collector, the data collector electrical signal is connected to the portable computer, data collection starts before the aircraft arrives above the measurement area, data collection stops after the sonic boom signal is successfully measured, and the measured sonic boom signal is analyzed and processed within the data collection time period. The invention can realize accurate measurement of ground sonic boom signals of supersonic aircraft.

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 invention relates to a seven-stringed plucked instrument desk sound amplification system and a seven-stringed plucked instrument desk, the sound amplification system comprises a microphone pickup circuit, a balance conversion circuit, a pre-amplification circuit, a post-stage power amplification circuit and a circuit structure with a small magnification step-by-step amplification, after a microphone collects an electric signal, filtering and noise reduction are carried out, then first-stage small magnification amplification is carried out, the amplification rate is 10 times, and the first-stage small magnification amplification is carried out; it is ensured that low level is not lost, high level is not overflowed, and effective collection of the whole range is ensured. After the signal is buffered, the signal is amplified at a small magnification of 10 times, so that a certain thrust is applied to the rear stage under the condition of ensuring the integrity of the signal. The amplification factor of 4.7 multiplying power is adopted in the last stage of amplification, so that the amplification factor of 470 times is adopted in the preceding stage in total, the output voltage amplitude can be effectively improved without distortion, and finally electro-acoustic conversion is carried out through the power amplifier circuit so as to achieve the purpose of ancient sound amplification. According to the invention, the audio frequency is expanded, frame loss or sonic boom of the audio frequency is avoided, and teaching training and large-scale playing are facilitated.

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.

A test system and method for simulating sonic boom waves based on electric sparks belongs to the technical field of ground simulation tests of sonic boom waves. It solves the problem in the prior art that sonic boom models do not simulate the atmospheric turbulence environment. By adjusting the turbulence generation system, a disturbance source is added to the flow field to generate a plane jet, the turbulence intensity is measured using a hot wire measurement system, and an N-type wave signal used to characterize typical sonic boom characteristics is generated with the help of an electric spark sonic boom source system. The sonic boom wave measurement system is used to measure the sonic boom wave signal and store the measurement data. The present invention is a ground test simulation system and method that can complete N-type waves, atmospheric turbulence, and the influence between the two of far-field sonic booms on the ground. It can take into account the influence of atmospheric turbulence on sonic booms, provide a ground simulation platform for the study of the interaction between sonic boom waves and atmospheric turbulence, and improve the prediction accuracy of far-field sonic booms.

The present application discloses a method, device, medium and product for predicting the sonic boom intensity of an aircraft, relating to the technical field of sonic boom prediction. The method obtains the surface pressure distribution data of the aircraft under different flight states based on the number and arrangement positions of the pressure measurement holes on the aircraft surface; obtains a data set of historical surface pressure distribution data and the corresponding ground sonic boom intensity of the historical surface pressure distribution data; trains a sonic boom prediction model based on the data set to determine a trained sonic boom prediction model; and inputs the surface pressure distribution data to be detected obtained in real time under the current flight state into the trained sonic boom prediction model to determine the ground sonic boom intensity corresponding to the surface pressure distribution data to be detected. The present application can predict the ground sonic boom intensity in real time and shorten the prediction time of the sonic boom intensity.

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.