165 results about "AxSTREAM" patented technology

This is an improvement of the 1968 Trailing Rotor V / STOL aircraft (Ref. 1). Rotors are mounted on wing-tip pods which can be tilted from the vertical to the horizontal aft position. Rotors are then stopped in the axial-flow condition and indexed to an azimuth position, aft of the wing trailing-edge. Rotor blades are then folded forward (blade-tips upstream of rotor-hubs) and locked into grooves in the tip-pods.The main improvements over Ref. 1 are: a smaller shift of center-of-gravity during transition to cruise mode, and an easier task of locking blades into the tip-pods.The main feature of the autorotative aft rotor tilt is that a soft-inplane rotor can be used, which reduces rotor weight. The blade-folding axis is also the blade-flapping axis.The autorotative mode is used frequently by helicopters during descent. It has been found to be a good, stable mode with rpm-stability. During aft tilt on the UFR, the rotors provide pitching stability to the airframe. During stopping of the rotor, the stability of the autorotating rotor eases the task of passing thru rotor resonance.Also, the wing can be swept back, which is desirable for 400 kt. cruise speeds and can be used for external stores such as fuel and weapons.

The present invention provides a nonlinearly stacked low noise turbofan stator vane. The stator is in an axial fan or compressor turbomachinery stage that is comprised of a collection of vanes whose highly three-dimensional shape is selected to reduce rotor-stator and rotor-strut interaction noise while maintaining the aerodynamic and mechanical performance of the vane. The nonlinearly stacked low noise turbofan stator vane reduces noise associated with the fan stage of turbomachinery to improve environmental compatibility. The stator vane has a characteristic curve that is characterized by a nonlinear sweep and a nonlinear lean.

The invention discloses an aerodynamic design method of a high-performance ship gas turbine compressor, aiming to solve the problems of complex pneumatic design operation and small application range of a ship gas turbine compressor based on ship power in China in the prior art. The aerodynamic design method of the high-performance ship gas turbine compressor comprises the steps: according to design requirements and targets and the overall design requirements of the compressor, conducting one-dimensional inverse problem design of a multi-stage axial flow compressor, and on the basis of a one-dimensional design scheme, performing S2 through-flow inverse problem design; comparing and analyzing three-dimensional matching characteristics of the gas compressor by adopting a multi-stage gas compressor three-dimensional numerical simulation method, carrying out three-dimensional numerical simulation, and comparing and analyzing the three-dimensional numerical simulation with one-dimensional and S2 calculation results; and returning the obtained data to the S2 through-flow inverse problem design, and repeatedly iterating to complete the S2 through-flow inverse problem design and the blade modeling design. The aerodynamic design method can be applied to all fields of gas turbines, including military, energy, electric power, traffic, environmental protection and the like.

In order to provide a high-head, large-delivery turbopump adapted for excellent suction performance and foreign matter passability by having characteristics of an inducer, axial flow vanes, and mixed flow vanes imparted together to centrifugal vanes, a progressively diameter-increased suction casing rear part (9b) has arranged therein an impeller (2) configured with two to four rotary vanes (13) wound around a hub (12), to define rotary channels (CA) having a vane outlet channel width (b2) of 26% in proportion to a vane inlet outer circumference diameter (d10), the rotary vanes being each respectively configured as a collision-less connection of an upstream axial-flow screw part (13a) provided with an inducer part (13a1) extending into a suction fluid path of a suction casing front part (9a) at a vane inlet angle (beta1) of 14°, an intermediate mixed-flow screw part (13b), and a downstream centrifugal screw part (13c).

The invention discloses an axial flow compressor stall surge prediction method based on deep learning, and belongs to the technical field of aero-engine modeling and simulation. The method comprises:firstly, preprocessing aircraft engine surge data, and dividing a test data set and a training data set from experimental data; secondly, sequentially constructing an LR branch network module, a WaveNet branch network module and an LR-WaveNet prediction model; and finally, performing real-time prediction on the test data: firstly, preprocessing the test set data in the same way, and adjusting thedata dimension according to the input requirement of the LR-WaveNet prediction model; according to the time sequence, by adopting an LR-WaveNet prediction model, giving the surge prediction probability of each sample; and by adopting an LR-WaveNet prediction mode, giving the surge probability of data with noise points along with time, and testng the anti-interference performance of the model. According to the method, the time domain statistical characteristics and the change trend are integrated, the prediction precision is improved, and certain anti-interference performance is achieved; and the active control performance of the engine can be improved, and certain universality is achieved.

The invention discloses an axial-flow fan and relative application, especially a new fan blade structure for reducing the fan noise, based on increasing wind amount, to meet the demands of air conditioner and improve the heat exchange efficiency. Wherein, said axial-flow fan comprises a central hub as the rotational axle and several wings radial arranged on the circumference of hub; the back edge of wing has an expansion part that liner expanding backwards the rotational speed of wing; said expanding part is formed by expanding from the inner edge at one side of hub to the outer edge, but not concave inwards to the rotational speed; the wing surface that facing the air-in side is a smooth plate; the diameter of said wing is between 400mm and 500mm; and its operation speed is controlled at 600-800rpm.

The invention relates to a pumping current blower fan with low noise and high wind rate, which belongs to non-positive-displace pump technical field. The shape of fan blade is back-swept, and its sectional plane is airfoil. The especial is that central axis is the center of rotation of the vane, it is cut by different radius cylinder face, geometry of the vane is fixed on by the numerical value of the chord An which is from the vane leading edge to its slurry radix axle wire, chord is spread by the sectional plane, X and Y in the sectional plane airfoil developed pattern. The advantage of the invention is big air output, low noise, high static pressure, and light weight.

This invention refers to the method of modeling and calculating aerodynamic characteristics of a multi-stage axial compressor using commercial software, method to identify the stable working range of module. The method includes step 1: object modeling; step 2: constructing the calculation model; step 3: problem solving. Step 4: results analysis.

The invention discloses a method for predicting surge frequency of an axial flow compressor. The method is used for solving the technical problem of poor practicability of the existing method for measuring the surge frequency of the axial flow compressor. The technical scheme is as follows: firstly, the experimental table of the axial flow compressor is physically and geometrically modeled to obtain the conservation control equation of the corresponding model, and after the numerical method is used to solve, the numerical result under the working condition of over-stall is analyzed by Fouriertransform, so as to obtain the surge frequency of the axial flow compressor. Because the simulation model of the axial flow compressor system is established, the construction of the experimental platform and the data acquisition system is avoided, the experimental data measurement process is transformed into the parameter value obtained by numerical solution, and then the numerical result is directly used for data analysis and processing, and the Fourier transform is used to analyze the surge frequency of the compressor, which saves a lot of manpower and material costs, and has good practicability.

The invention relates to a combined surge eliminating method. The method comprises the steps that a first combined surge eliminating scheme and a second combined surge eliminating scheme are set according to the working condition of a gas compressor; when the rotating speed of the gas compressor is lower than the set rotating speed, the first combined surge eliminating scheme is executed until surge is eliminated; when the rotating speed of the air compressor is higher than the set rotating speed, the first combined surge eliminating scheme is executed firstly, the first combined surge eliminating scheme is stopped when surge is eliminated, and the second combined surge eliminating scheme is executed until surge is eliminated while the first combined surge eliminating scheme is kept to beexecuted when surge is not eliminated; and the first combined surge eliminating scheme comprises opening an exhaust throttle valve and an inter-stage air entraining control valve, and the second combined surge eliminating scheme comprises reducing the rotating speed of the air compressor and adjusting the angle of an adjustable blade. The technical problem that when the surge strength is large, surge is difficult to effectively eliminate at a time can be solved, the timeliness and effectiveness of surge eliminating can be ensured through different surge eliminating modes aiming at low-rotating-speed surge and high-rotating-speed surge in the air compressor, and the working safety of the air compressor is effectively improved.

The invention provides an adaptive attack angle method of an axial flow air compressor. The method comprises the following steps: by building a calculation formula, obtaining a relation between a rotor pitch diameter air feeding attack angle and an adjustable guide vane geometric outlet angle; then determining an adjustable value or an adjustable range of the adjustable guide vane geometric outlet angle based on the relation between the rotor pitch diameter air feeding attack angle and the adjustable guide vane geometric outlet angle according to acquired temperatures, total pressures and static pressures of all positions on the section of an adjustable guide vane inlet and a pre-determined a target value or a target range of the rotor pitch diameter air feeding attack angle; and adjusting a rotating angle of an adjustable guide vane according to the obtained adjustable value or adjustable range of the adjustable guide vane geometric outlet angle, thus realizing an adaptive attack angle of the axial flow air compressor and enabling the attack angle of a rotor of the axial flow air compressor to be maintained in a proper range. Therefore, the axial flow air compressor can keep a good working state within a wide working range, and the workload and the working difficulty of an operator can be reduced.

The invention aims to provide a multi-dimensional collaborative design method for a multistage axial flow compressor of a ship gas turbine, and the method comprises the following steps: repeatedly iterating links such as one-dimensional design and analysis, two-dimensional design, quasi-three-dimensional design, blade modeling design, quasi-three-dimensional analysis, full-three-dimensional analysis, collaborative design and the like; and obtaining the aerodynamic design scheme of the multistage axial flow compressor of the ship gas turbine meeting the design index requirements. The method hasthe characteristic of cross-dimension design, design parameters of different dimensions can be interactively used, only three-dimensional CFD calculation with few times is needed to initialize the loss model, and after the initialization work of the loss model is completed, the loss and performance parameter information of the compressor can be obtained in a quasi-three-dimensional design stage without three-dimensional CFD calculation, so that the design efficiency of the compressor is improved. Therefore, a high-quality compressor aerodynamic design scheme can be quickly completed, a largeamount of calculation time is saved, the design period is effectively shortened, and the method is very suitable for engineering design application.

The invention discloses a blade, a modeling method and manufacturing method thereof and an air compressor. The modeling method comprises the steps that a flowing channel is sequentially divided into an inlet semi-vaneless area, a channel expansion area and an outlet semi-vaneless area in the gas flowing direction; front edge side end points of a suction surface molded line and a pressure surface molded line of the flowing channel are determined; blade molded lines forming the boundaries of the inlet semi-vaneless area, the channel expansion area and the outlet semi-vaneless area are sequentially determined, and the channel expansion angle of the whole flowing channel is regulated and controlled; on the basis of the determined molded lines, modeling of the primitive-level blade profiles is completed, and then the primitive-level blade profiles are stacked to form the three-dimensional blade. The modeling method has the advantage that the diffusion degree of an expansion channel is controllable, the quality of flowing airflow in the expansion channel of the axial flow air compressor blade is improved helpfully, pneumatic matching between blade rows is improved, and the efficiency of the axial flow air compressor is improved; and the modeling method has good applicability to radial / axial diffuser blades of a centrifugal air compressor and high-low pressure stage rotor / stator blades of a turbine.

The invention belongs to the field of compressor performance simulation, and relates to a stage characteristic matching method of a multi-stage axial flow compressor, which comprises the following steps: selecting one-dimensional calculation parameters of stage characteristics; obtaining a stage characteristic one-dimensional calculation parameter value in the inlet direction and the middle diameter position of each stage of rotor blade of the multistage axial flow compressor; defining the stage characteristic one-dimensional calculation parameter value of the first stage of the multi-stage axial flow compressor as an inlet condition, and defining the stage characteristic one-dimensional calculation parameter values of other stages except the first stage as inlet boundary conditions; single-stage characteristic values of other stages except the first stage are calculated, and the single-stage characteristic value of the first stage and the joint calculation stage characteristic value of the multi-stage axial flow compressor are calculated; drawing a stage characteristic analysis curve of each stage of the multi-stage axial flow compressor by using the joint calculation stage characteristic value and the single-stage characteristic value; and inter-stage parameter matching of the multi-stage axial flow compressor is judged. According to the stage characteristic matching method designed by the invention, stage characteristic matching in the design scheme of the multi-stage axial flow compressor can be quickly analyzed, and the method plays an important role in judging the matching performance of the inter-stage parameters of the compressor.

The invention relates to a method for predicting characteristics of a gas compressor according to design indexes, and belongs to the field of gas compressor characteristic prediction. The method comprises prediction of characteristic parameters of peak efficiency points of all the isotactic speed lines and prediction of continuation of all the isotactic speed lines towards a near surge area and a near blockage area. Data input in the prediction process is only compressor design indexes (design point pressure ratio and efficiency), and compressor blade profile data are not needed. According to the design point pressure ratio, three coefficients are constructed, firstly, the peak efficiency point characteristics of each isorotation speed line are predicted, and then characteristic lines are expanded to a surge boundary and a blockage boundary respectively. The formulas and parameters used in the method are derived from mining of historical gas compressor characteristic data. The prediction method provided by the invention can be used for predicting the performance of an axial flow compressor and a centrifugal compressor, is particularly suitable for predicting the performance of the compressor when no blade profile data exists in the initial stage of the overall design of a gas turbine and an aero-engine, and is simple and convenient in calculation and relatively high in precision.

The invention discloses a method for predicting dynamic stall process of an aviation axial flow compressor under complex air inlet distortion. The method comprises the steps that a blade wall fixing effect is described through a distributed force source, and a dynamic relation between a local blade element runner force source and an airflow parameter is effectively constructed based on a loss coefficient and a backward angle concept. The method has the capability of quantitatively describing the dynamic stall evolution process of the multi-stage axial flow compressor under complex air inlet distortion. Compared with a traditional two-dimensional calculation method and a high-order three-dimensional CFD calculation method, the method has the advantages that the calculation efficiency can be considered, and meanwhile, the aerodynamic instability characteristics of the aviation axial flow compressor under complex air inlet distortion disturbance including rotational flow and radial / circumferential space combination distortion characteristics are effectively reflected, and a technical support is provided for high-fidelity aerodynamic stability design of an advanced aero-engine.

The invention aims to provide a multistage axial flow compressor aerodynamic design method based on key dimensionless load control parameters, which comprises the following steps of: repeatedly iterating links such as one-dimensional inverse problem design and characteristic analysis, S2 inverse problem design, blade modeling design, three-dimensional CFD analysis and the like; and obtaining a multi-stage axial flow compressor aerodynamic design scheme meeting design index requirements. Customized design of the internal load of the multi-stage axial flow compressor is achieved through three dimensionless parameters including the flow coefficient, the load coefficient and the reactivity, and the problem of space load matching between design points and low working conditions in pneumatic design of the multi-stage axial flow compressor is effectively solved; and the pneumatic design of the multistage axial flow compressor achieves the parameterization and refinement degree, the design precision is improved, and the design period is shortened. Meanwhile, the method is not limited to a multi-stage axial flow compressor of a gas turbine, but also suitable for the pneumatic design process of various industrial axial flow compressors and aero-engine axial flow compressors / fans.

The invention provides an axial flow compressor rotation stall prediction method based on a stacked long short-term memory network, and belongs to the technical field of aero-engine modeling and simulation. The method comprises the following steps: firstly, selecting and preprocessing surge experiment data of a certain type of aero-engine, and dividing the data into a training set and a test set; secondly, building and training a Stacked LSTM model, carrying out the real-time prediction on a test set by utilizing the finally trained model, and giving model loss and evaluation indexes; finally, carrying out real-time prediction on the test data by adopting a Stacked LSTM prediction model, and giving a change trend of the surge probability along with time according to a time sequence. According to the method, the time domain statistical characteristics and the change trend are integrated, and the prediction precision is improved; and the active control performance of the engine can be improved, and certain universality is achieved.

The invention belongs to the technical field of turbomachinery, and discloses a framework forming axial flow turbine blade design method, which decomposes blade design into a mean camber line design, a thickness distribution design and a stacking line design. According to the mean camber line design, two control parameters are set on a mean camber line to represent the mean camber line, and according to the thickness distribution design, two control parameters are set on thickness distribution to represent thickness distribution and used for adjusting the blade profile. The stacking line design is that the stacking line is expressed in a parameter form and is used for controlling the radial curvature of the blade to generate a straight blade and a bent and twisted blade; the characteristic section formed by the mean camber line and the thickness distribution is arranged according to a bending and twisting line formed by the radial installation angle and the stacking line, and the adjustable three-dimensional blade is generated. Through eight control parameters, an efficient bending and twisting combined axial flow turbine can be quickly generated through parametric modeling. The invention is wide in applicability, can be expanded to other turbomachinery, and has a great application prospect.

The invention aims to provide a method for designing a spoiler type processing casing of a return cavity of a gas compressor of a ship gas turbine, which comprises the following steps of: selecting a typical working condition according to a pneumatic design result of the gas compressor, and determining an axial characteristic position of the spoiler type processing casing of the return cavity; extracting an inlet absolute airflow angle and an inlet relative airflow angle at the blade top of the movable blade at the feature position so as to determine an air inlet radial angle and an air outlet radial angle of the spoiler; selecting three influence factors including the outer diameter of the spoiler, the length of the spoiler and the number of the spoilers to design a structural scheme; respectively carrying out numerical calculation on different structural design schemes to determine an optimal scheme; and performing performance analysis on the optimal scheme under the design working condition, and judging whether the design point performance of the scheme reaches the standard or not. The method is not only limited to a ship gas turbine axial flow compressor, but also suitable for various industrial gas turbine axial flow compressors with processing casings and aero-engine axial flow compressors.

The invention provides a modeling method for a non-axisymmetric end wall of an axial flow impeller, which comprises the following steps: for end wall molded surfaces with different concavities and convexity, firstly, dividing a configuration area, and flexibly constructing various complex end wall structures on the basis of applying fewer variables by adopting a circumferential three-order Fourierseries curve; and then reading a three-order Fourier series curve in the circumferential direction of the blade, enabling spline curves to subjected to smooth connection in the axial direction, and forming a non-axisymmetric end wall model. A three-order Fourier series curve mode is adopted in the circumferential direction to accord with the distribution rule of a blade channel pressure field, the axial section area of a flow channel can be kept almost unchanged, the total loss near the end wall is reduced, flowing separation of airflow is reduced, the outlet flow field and the airflow angleare more uniform, and the performance of a fan and a compressor blade is improved.

The invention aims to provide a method for designing a grooved ring type processing casing of a ship gas turbine compressor, which comprises the following steps of: selecting a typical working condition, and obtaining an axial characteristic position of the grooved ring type processing casing by analyzing pneumatic parameter distribution of a flow field close to a surge point under the typical working condition; extracting an inlet absolute airflow angle and an inlet relative airflow angle at the top of the moving blade at the axial characteristic position close to the surge point under a typical working condition so as to determine a slotting angle of the grooved ring; determining the lengths of the chamber 1 and the chamber 2 of the grooved ring processing casing; selecting three influence factors including the slotting radius, the slotting length and the slotting number of the slotted ring to design a structural scheme; performing numerical calculation on different structural design schemes to determine an optimal scheme; and performing performance analysis on the optimal scheme under the design working condition. The method is not only limited to a ship gas turbine axial flow compressor, but also suitable for the processing casing design process of various industrial gas turbine axial flow compressors with processing casings and aero-engine axial flow compressors.

The invention provides a method and device for estimating aerodynamic performance of a multistage axial flow compressor. The method comprises the following steps that first characteristic line families and first complete machine characteristic lines of all stages in the design state of the compressor are obtained in advance, and each first characteristic line family comprises a plurality of firstcharacteristic lines under the conversion rotating speed condition; the reserved-stage characteristic line families in the state, deviating from the design state, of hardware in the compressor are obtained, and each characteristic line family comprises a plurality of characteristic lines at a conversion rotating speed; the second complete machine characteristic line in the state, deviating from the design state, of the hardware in the compressor is obtained; and the first complete machine characteristic line is compared with the second complete machine characteristic line to obtain the influence of deviating from the design state of the hardware. The method and the device provided by the invention have higher estimation efficiency and higher estimation accuracy.

The invention provides a method for realizing aeroelasticity optimization design of turbomachinery by considering random detuning. Based on flutter calculation of an influence coefficient method and forced vibration analysis of a basic detuning model and a Monte Carlo method, on the premise that aerodynamic performance and static stress are restrained, industrially-machined blades have similar vibration characteristics in different installation modes. Flutter stability is remarkably improved, and blade forced vibration response under condition of random detuning is reduced. According to the method, for the selected axial flow fan blade, on the premise that aerodynamic performance and static stress are restrained, the minimum aerodynamic damping of the selected axial flow fan blade is increased by about 19%, and the maximum forced vibration amplitude under random detuning is reduced by about 18%.

The invention aims to provide a method for designing a low-working-condition rotatable guide vane corner rule of an axial flow compressor of a ship gas turbine, which comprises the following steps: extracting pneumatic parameters at the characteristic section positions of movable vanes in the adjacent rows of rotatable guide vanes of each column of the compressor, decompressing the low-working-condition reduced rotation speed and the design point reduced rotation speed of the compressor, and calculating the circumferential speed of the movable vanes; setting a reduced flow value, and calculating the axial speed of an inlet of each column of moving vanes; and giving an expected inlet attack angle of the next-stage moving vane, obtaining an inlet relative airflow angle of the next-stage moving vane after each column of rotatable guide vanes rotate, solving a rotating angle of each column of rotatable guide vanes, and obtaining a rotating angle rule of the rotatable guide vanes of the gascompressor at different folding rotating speeds. By means of the invention, the ideal multi-column rotatable guide vane combined regulation and control corner rule can be rapidly obtained, resource and time consumption in the design process is effectively reduced, the workload of designers is reduced, the surge margin index of the air compressor under the low working condition can be effectivelyimproved, and the invention is very suitable for engineering design application.

The invention proposes a splitter lip 22 for a low-pressure compressor of an axial turbomachine for an aeroplane. The splitter lip 22 comprises an upstream annular wall 36 made of metal with a circular leading edge 30, and a downstream annular partition 38 made of an organic-matrix, short-fibre composite material. The splitter lip also supports an outer shroud 28 for a stator upstream of the compressor. The upstream wall comprises an annular anchoring portion 40 arranged in the thickness of the downstream partition 38 so as to anchor the partition and the wall to one another. The anchoring portion 40 has distributed hexagonal openings in order to increase anchoring. The invention also proposes a method for producing, by moulding, a bi-material splitter lip or a mixed lip.

The invention discloses an S2 flow field diagnosis method based on inter-stage measurement parameters of an axial flow compressor. The S2 flow field diagnosis method specifically comprises the steps of 1, establishing a corresponding S2 flow field calculation streamline curvature control equation according to the inter-stage measurement parameters of the multi-stage axial flow compressor; 2, according to the fact that the possible total temperature rise of the multi-stage compressor is high, considering the influence of inter-stage measurement parameters by adopting a variable specific heat working medium; 3, obtaining stator lag angle and entropy distribution parameters in the flow field according to a thermodynamic relationship and a stator model; 4, performing flow field calculation toobtain an axial flow compressor S2 flow field inversion result according to the inter-stage measurement parameters; and 5, performing analysis and diagnosis according to calculation results such as flow field parameters and performance parameters obtained by inversion. According to the method, the variable specific heat working medium is adopted, and the accuracy is higher when the conditions thatthe stage number of the compressor is large and the temperature rise is large are treated; and the assumption that the static pressure of measuring points at different heights is equal to the wall surface static pressure in the multi-stage compressor test is broken through, the flow field result automatically meets the radial balance equation, the S2 flow field diagnosis precision is higher, andthe compressor design cycle is reduced.

The invention aims to provide an axial flow compressor S2 inverse problem design method based on dimensionless load control parameters. An S2 aerodynamic design scheme of an axial flow compressor is obtained through spanwise distortion rule design, spanwise loss design and typical S2 flow surface streamline curvature method solving. Customized design of prewhirling of all stages of moving blade inlets of the axial flow compressor is conducted through two key dimensionless parameters including the load coefficient and the reaction degree, active control over the space load of the compressor in the S2 design stage is achieved, and the S2 inverse problem through-flow design of the axial flow compressor reaches the parameterization and refinement degree; the pneumatic design precision of the axial flow compressor is effectively improved, the pneumatic performance is improved, and the method is very suitable for engineering design application. Meanwhile, the invention is not limited to a gas turbine axial flow compressor, but also suitable for the pneumatic design process of various industrial axial flow compressors and aero-engine axial flow compressors / fans.