17 results about "Gas dynamics" patented technology

The invention provides a predictive maintenance method and system for an air suspension fan based on digital twinning, and the method comprises the steps: collecting an air flow abrupt change characteristic when high-temperature dust-containing waste gas flows into the air suspension fan, transmitting the air flow abrupt change characteristic to a digital twinning system, and determining the air flow abrupt change characteristic; extracting a local clearance reduction trend of the foil and the bearing from the non-uniform expansion amount of the foil, and calculating space distribution after clearance reduction to obtain a target gas film thickness value; the target gas film thickness value is compared with a preset collision and abrasion threshold value, and if the target gas film thickness value is smaller than the threshold value, the collision and abrasion risk probability is evaluated through a Monte Carlo method; if the collision and abrasion risk probability exceeds a preset safety range, calculating a gas film rigidity fluctuation range based on a gas dynamics theory; and extracting a vibration amplitude analysis result from the dynamic response parameters of the foil bearing of the air suspension fan, and calculating a stability prediction index.

The invention relates to the technical field of high-rise building elevators, and provides a high-rise elevator emergency air pressure slow descending device which comprises an elevator car, first wedge-shaped blocks fixed to the two sides of the elevator car, a plurality of slow descending mechanisms embedded in the two sides of an elevator shaft in a gradient mode and an air pressure bottom sinking mechanism arranged at the bottom of the elevator shaft. The multi-layer buffering mechanism combining non-Newtonian fluid damping, pneumatic pressurization and elastic supporting has the synergistic effect, the shear thickening effect of non-Newtonian fluid in the box body is activated through the stirring components (the spiral pressing pieces and the stirring blades), the fluid viscosity is greatly improved instantly when the lift car descends suddenly, dynamic damping force is formed, and the dynamic damping effect is achieved. The first air cylinder is used for compressing air to drive the air bag to expand, the reverse thrust assembly (the second air cylinder and the lower pressing plate) is combined to achieve conversion from air pressure energy to mechanical energy, secondary buffering is formed, non-linear and self-adaptive energy absorption is achieved through coupling of fluid mechanics (non-Newtonian fluid) and gas dynamics, and the device is particularly suitable for high-speed falling scenes.

The invention relates to the technical field of aerospace engines, in particular to an axial expansion combustion chamber and a rotary detonation engine, which comprise a base, a gas distribution assembly and an axial expansion combustion chamber which are coaxially and sequentially arranged along the axial direction, the axial expansion combustion chamber is improved aiming at the technical defect that the flow speed of supersonic airflow in an existing uniform-section combustion chamber is reduced due to the gas dynamic effect and the wall surface heat sink effect, so that the inner wall of the axial expansion combustion chamber expands outwards in the axial direction of the airflow; therefore, the sectional area of an inner channel of the axial expansion combustion chamber is gradually expanded in the axial direction, and an expansion type flow channel is formed. Compared with the prior art, the inner channel has the beneficial effects that through the physical geometric structure change of axial expansion of the inner channel, the total pressure loss and speed attenuation of airflow in the flowing process are successfully compensated, the energy release efficiency in the combustion chamber is remarkably improved, and the self-sustaining capacity of an engine in the combustion process is effectively enhanced.

The application belongs to the technical field of aero-engines, and particularly relates to a method for evaluating a Mach number and total static pressure at an inlet of a main combustion chamber, comprising the following steps: S1: arranging a test rake at a windward surface of a support plate of a pre-expander of the main combustion chamber, and measuring and obtaining total pressure P31 and total temperature T31 of a support plate section; S2: obtaining known physical parameters of an outlet section of a high-pressure compressor, including a flow area A3 and an air mass flow W3 at the inlet of the main combustion chamber; S3: based on the total pressure P31 and the total temperature T31 of the support plate section and the parameters A3 and W3 of the outlet section of the compressor, a Mach number Ma3 of the outlet section of the compressor is obtained by iteratively solving a gas dynamics relationship; and S4: based on the obtained Mach number Ma3, total pressure P3 and static pressure Ps3 of the outlet section of the compressor are calculated.

This invention discloses a semi-analytical method and system for calculating the stiffness and damping coefficients of foil gas dynamic bearings, relating to the field of air conditioning and clothing temperature control technology. It aims to address the shortcomings of traditional Reynolds equation methods in terms of accuracy and universality, as well as the low computational efficiency and mesh distortion issues of existing CFD methods. The method includes: constructing a fluid-structure interaction (FSI) computational model to calculate and obtain the static equilibrium position of the journal; applying an eccentricity perturbation at the static equilibrium position and performing an additional steady-state FSI calculation to obtain the four stiffness coefficients of the foil bearing using a first set of analytical formulas; applying small-amplitude simple harmonic vibrations along the x and y directions to the journal and performing two transient FSI calculations to obtain the four damping coefficients of the foil bearing using a second set of analytical formulas. This invention combines high accuracy and high computational efficiency, providing an efficient and reliable analytical tool for evaluating the dynamic characteristics of foil bearings and designing the stability of the rotor systems they support.

The invention provides a method and a system for constructing a thermal hydroelastic equivalent model, and the method comprises the steps: obtaining the structural dynamic characteristics of a real aircraft in a target thermal environment through thermal-fluid-solid multi-physics field coupling simulation, and taking the structural dynamic characteristics as a target value of equivalent design; decoupling the aerodynamic thermal environment influence represented by the target value, and equivalently converting the aerodynamic thermal environment influence into design requirements for physical attributes of an equivalent model at normal temperature; based on the design requirements, an equivalent model with the structural dynamic characteristics matched with the target value under the normal temperature condition is designed through inversion of an optimization method. According to the design method of the thermodynamic elasticity model of the equivalent temperature effect, the influence of the temperature on the structure is equivalent to weakening of the structural strength, and related research on the thermodynamic elasticity can be carried out in a conventional hypersonic wind tunnel.

The application provides an air suspension fan predictive maintenance method and system based on digital twinning, comprising: collecting airflow mutation characteristics when high-temperature dust-containing waste gas rushes into the air suspension fan, transmitting to a digital twinning system, and determining the airflow mutation characteristics; extracting a foil and bearing local gap narrowing trend from a foil non-uniform expansion amount, calculating a spatial distribution after the gap is narrowed, and obtaining a target gas film thickness value; comparing the target gas film thickness value with a preset collision and grinding threshold value, if the target gas film thickness value is less than the threshold value, then evaluating a collision and grinding risk probability through a Monte Carlo method; if the collision and grinding risk probability exceeds a preset safety range, then calculating a gas film stiffness fluctuation range based on a gas dynamics theory; and extracting vibration amplitude analysis results from air suspension fan foil bearing dynamic response parameters, and calculating a stability prediction index.

The invention provides a gas flow control method and system applied to a solid-state working medium electric propulsion system, and the method comprises the steps: firstly constructing a second-order nonlinear prediction model from heating power to output flow in order to solve the problems of strong nonlinearity, multivariable coupling and strict safety constraints in a solid-state iodine sublimation process; the model accurately depicts a complete physical process of heating sheet thermal inertia, contact thermal resistance, iodine phase change dynamics and pipeline gas dynamics. And on the basis, rolling optimization is executed in each control cycle: a future flow trajectory is predicted based on a current state and a model, and an optimization problem with constraints is solved online to obtain an optimal heating power sequence. According to the method, high-precision and overshoot-free tracking of the flow is achieved, it is strictly guaranteed that the system operates within safety constraints, and the performance and reliability of the solid-state electric propulsion system are remarkably improved.

The application discloses a kind of aircraft model rare gas dynamics numerical simulation method, it is related to numerical simulation field, the method includes: step 1: based on aircraft model parameter construction tetrahedron grid file;Step 2: the tetrahedron grid file is handled, obtains the tetrahedron grid file based on red-black tree;Step 3: setting inflow parameter, numerical simulation is carried out using the tetrahedron grid file based on red-black tree, obtains simulation result;The present application can effectively improve internal face element de-redundant process, improve the efficiency of tetrahedron grid geometry preprocessing program, to improve the efficiency of aircraft model rare gas dynamics numerical simulation method, compared with the processing mode based on array loop, computing efficiency is significantly improved.

The invention relates to the technical field of pneumatic turbines, in particular to a rotating wheel for a pneumatic turbine, and solves the problem that an existing rotating wheel of the pneumatic turbine is low in efficiency. The rotating wheel comprises a rotating wheel hub, a rotating wheel outer ring and a plurality of wing-shaped blades, and the wing-shaped blades are connected between the hub and the outer ring to form a divergent blade runner; the wing-shaped paddle is in a three-dimensional twisted wing shape, and the wing inner side and the wing outer side of the wing-shaped paddle have different relative thicknesses, relative cambers and mounting angles; sub-runners are formed between the adjacent paddles, the cross sectional areas of the sub-runners are gradually reduced from an inlet to an outlet, and the airflow directions of the inlet and the outlet are configured to be axially parallel or basically parallel to the rotating wheel. The rotating wheel can be used in multi-stage series connection, and parameters of each stage of blade are gradually changed and optimized in the airflow direction. Through the collaborative design of the divergent flow channel and the three-dimensional twisted blades, the full-flow-channel airflow organization is optimized, the energy conversion efficiency is remarkably improved, the overall efficiency up to 89.8% is achieved in a four-stage pneumatic turbine, and the structural universality is high.

This invention relates to the field of helicopter and rotorcraft unmanned aerial vehicle (UAV) technology, specifically to a method for estimating the jet thrust of a jet rotor, comprising: obtaining first temperature and pressure parameters of the aero-engine outlet gas based on the characteristic parameters of the compressor, combustion chamber, and turbine of the aero-engine; calculating heat transfer losses based on the heat transfer area, geometric parameters, and temperature parameters of the main shaft duct; obtaining second temperature and pressure parameters of the outlet gas of the main shaft duct based on the heat transfer losses and the pressure loss at the first bend; dividing the rotor duct into multiple segments and calculating the intra-segment losses in each segment; obtaining third temperature and pressure parameters of the gas before the nozzle based on the pressure loss at the second bend; and inputting the optimal nozzle cross-section, the third temperature, and the pressure parameters into a gas dynamics model for processing to obtain the jet thrust. This invention can improve the accuracy of jet thrust calculation for jet rotors.

The application discloses a kind of hypersonic vehicle global active thermal protection methods, it is related to hypersonic vehicle thermal protection field.Gas dynamic heat peak value is greater than given threshold value, jet port opens, cooling working medium is injected flow field from jet port along the free flow reverse direction, and forms backflow area and covers in aircraft head, realizes head area thermal protection;Cooling working medium reconfigures the shock wave of coming flow, makes the arc shock wave of aircraft head become conical shock wave, and forms reattachment shock wave in aircraft shoulder area;Magnet is installed in aircraft inner wall surface, generates magnetic field and acts on plasma in flow field and generates external thrust force, pushes reattachment shock wave away from wall surface, guides cooling working medium to downstream, realizes global thermal protection.The effect and benefit of the application are: cooling working medium demand is small, thermal protection area coverage area is large, does not change aircraft shape structure, and reliability is high.

This invention relates to the field of fluid mechanics, disclosing a method, apparatus, device, and storage medium for determining instability in rarefied gas flows. By constructing a physical control model based on gas kinetic theory and solving the baseline flow field, modal and nonmodal analyses are performed on the disturbance evolution model. On one hand, eigenvalues ​​of the disturbance growth rate are obtained to determine the absolute instability risk. On the other hand, singular value decomposition is used to obtain the maximum energy growth factor of the initial disturbance within a finite time to quantify the transient energy growth effect caused by the nonorthogonality of linear operators. Finally, the flow instability risk is output based on the combined determination result of the eigenvalues ​​and the maximum energy growth factor. This invention solves the technical problem that existing kinetic modal analysis methods cannot identify subcritical instability due to neglecting the nonorthogonality of linear operators, achieving comprehensive identification of both absolute and subcritical instability, and significantly improving the accuracy and reliability of rarefied gas flow instability prediction.

The application discloses a method for estimating liquid nitrogen flow of a low-temperature wind tunnel test state, belongs to the field of experimental aerodynamics, and aims to solve the problem that the proportional-integral-derivative closed loop control is adopted to adjust system output in the current low-temperature wind tunnel operation, the space difference between the total temperature measurement position and the liquid nitrogen injection position is limited, and the mixing heat transfer process leads to serious hysteresis of the total temperature control, greatly increases the liquid nitrogen consumption, and reduces the test efficiency. In combination with the high-speed airflow convection heat exchange and gas dynamics, the application aims at the liquid nitrogen flow estimation problem of the low-temperature wind tunnel, proposes a calculation method of the liquid nitrogen flow based on the target values of the total pressure, the total temperature and the Mach number, and can calculate the required liquid nitrogen flow according to the test state before the test, greatly improves the low-temperature wind tunnel test efficiency, and reduces the operation cost. The application can be quickly applied to other similar wind tunnels, has good engineering applicability, high application value and good application prospect.

The application discloses a small-sized supersonic shock tunnel test section with easy assembly and disassembly, and belongs to the field of aerospace technology. The supersonic shock tunnel test section comprises a first flange connected with a shock tube, and a second flange connected with a wind tunnel; a round-to-square rigid section, a connecting fixed section and a test observation section are sequentially connected between the first flange and the second flange; and a nozzle is detachably installed in the test observation section. The fixed parts and the connecting parts of the test section are easy to assemble, the nozzle rigid section and the round-to-square section are easy to disassemble, a more efficient experiment process is realized, and compatibility of gas dynamics teaching practice and scientific research experiment is achieved; secondly, the longitudinal observation window and the side observation window are simultaneously used, optical conditions are more favorable for nozzle starting and working process, optical observation means such as the schlieren method, the fluorescent oil flow method and the sheet light tomography method can be applied to analyze the flow field mechanism, and the supersonic shock tunnel test section provides an equipment basis for cold flow experiment of an advanced nozzle of a solid rocket engine.

The utility model relates to the technical field of nuclear fusion and discloses an aerodynamics trap fusion system, the aerodynamics trap fusion system is provided with an energy recovery device used for the aerodynamics trap fusion system, the energy recovery device comprises a magnetic field generation assembly and a power generation assembly, the magnetic field generation assembly is arranged on an aerodynamics trap, and the power generation assembly is arranged on the aerodynamics trap. The magnetic field generation assembly can generate a deflection magnetic field and can change the movement direction of the plasma escaping from the gas dynamic trap; the power generation assembly is arranged on the gas dynamic trap, the power generation assembly can receive the plasma acted by the magnetic field generation assembly so as to form a potential difference capable of outputting current, and the gas dynamic trap fusion system can solve the technical problem of how to avoid waste caused by plasma dissipation loss. The energy recovery device in the gas dynamic trap fusion system can recycle escaped plasmas, avoid energy waste of the plasmas and improve the utilization rate of the plasmas in the gas dynamic trap.

The invention discloses an accelerated analysis method for aerodynamic characteristics of a free molecular flow of an aircraft, belongs to the field of numerical simulation of rarefied gas dynamics, and aims to solve the problem of low shadow analysis calculation efficiency of a ray tracing surface element method in aerodynamic analysis of an extremely low earth orbit aircraft. The method comprises the following steps of: initializing parameters and reading a surface grid; establishing a uniform background grid as a spatial index, and quickly associating surface triangular surface elements to corresponding sub-grids; an OpenMP parallel instruction is adopted, and ray tracing-based shielding judgment is executed on a large number of surface elements at the same time; and calculating the aerodynamic contribution of the unshielded surface element according to the free molecular flow theory, and integrating to obtain the overall aerodynamic force. According to the method, the calculation complexity of shadow analysis is remarkably reduced from O (N) through cooperation of spatial index and parallel calculation, and efficient and high-precision prediction of the aerodynamic characteristics of the extremely-low-orbit aircraft is achieved on the premise that precision is guaranteed.