CN113361211A - Method, system, equipment and medium for calculating aerodynamic stability of turboshaft engine - Google Patents

Method, system, equipment and medium for calculating aerodynamic stability of turboshaft engine Download PDF

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CN113361211A
CN113361211A CN202110405460.2A CN202110405460A CN113361211A CN 113361211 A CN113361211 A CN 113361211A CN 202110405460 A CN202110405460 A CN 202110405460A CN 113361211 A CN113361211 A CN 113361211A
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王召广
杨宇飞
王旭
屠宝锋
胡骏
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Hunan Aviation Powerplant Research Institute AECC
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Abstract

The invention discloses a method, a system, equipment and a medium for calculating the aerodynamic stability of a turboshaft engine. The calculation model adopted by the invention only focuses on parameters of the inlet and outlet sections of the compressor, the whole combined compressor can be modularized, and the static pressure, the static temperature, the mass flow and the circumferential speed are used as variables to establish the equation set to describe the flow in each control body, so that on one hand, all variables can be obtained through formula conversion of the equation set based on the four basic variables, and on the other hand, the four basic variables are convenient to obtain through a test means, and the data obtaining process is simplified.

Description

Method, system, equipment and medium for calculating aerodynamic stability of turboshaft engine
Technical Field
The invention relates to the technical field of aerodynamic stability evaluation of a turboshaft engine, in particular to a method, a system and equipment for calculating aerodynamic stability of the turboshaft engine and a computer readable storage medium.
Background
The helicopter has the advantages of vertical landing and hovering and the like which are not provided by fixed wing aircrafts, is limited by places and is convenient to use, and is widely applied to the military and civil fields, while the turboshaft engine has the advantages of large power-to-weight ratio, high reliability, good economy, long voyage, convenience in maintenance and the like, and is always the main power of the helicopter since the 20 th 60 s. When the turboshaft engine is applied to a helicopter, the helicopter needs to meet different taking-off and landing environments such as sand, grassland, mountainous areas, urban areas and the like, the flying height is not high, foreign matters are easy to adsorb, and the operation safety of the turboshaft engine is damaged by the suction of the foreign matters such as large-size sand grains, birds, tree leaves and the like, so that a protective cover is usually installed at the inlet of an air inlet channel. According to the current research conclusion, the protective cover can form certain pressure distortion due to the fact that the protective cover comprises a plurality of thicker supporting rods, in addition, the boundary layer of an air inlet channel can be separated when non-horizontal flight is carried out, pressure distortion can also be formed, and pressure distortion can also be caused by the downward washing air flow of a special rotor wing of a helicopter. Therefore, intake pressure distortion is a major factor causing a reduction in aerodynamic stability of the turboshaft engine. In the development and use process of the turboshaft engine, the problem of aerodynamic stability of the engine with aerodynamic instability of components or systems exists, the phenomenon is represented as rotating stall of a compressor or surge of the whole compression system, and the performance and stable working range of the turboshaft engine are limited by the aerodynamic stability. Therefore, the intake pressure distortion will affect the aerodynamic stability of the turboshaft engine, reduce the stability margin, and easily cause the engine to enter unstable working states of rotating stall, surging and the like, thereby causing the performance reduction of the engine and even serious operation safety.
Compared with a turbojet engine and a turbofan engine, the turboshaft engine has the remarkable characteristic that the turboshaft engine usually adopts a centrifugal compressor or an axial flow-centrifugal combined compressor. Foreign scholars find that the centrifugal compressor and the axial compressor have large difference in pneumatic instability through research, and although the centrifugal compressor and the axial compressor have many same flow phenomena, the centrifugal compressor and the axial compressor have difference in flow characteristics causing unstable flow states (rotating stall and surge), and particularly the difference is more obvious for the compressor with a high pressure ratio. From the current level of knowledge of the unsteady flow regime of the compression system, the basic problems and the basic research ideas concerning the aerodynamic stability of the centrifugal compressor and the axial compressor are the same in terms of overall system characteristics, but when specific flow phenomena inducing stall are specifically studied, the difference between the centrifugal compressor and the axial compressor will be manifested. Therefore, the existing calculation model of the turbojet and turbofan engine cannot be applied to the aerodynamic stability evaluation of the turboshaft engine. At present, no calculation model for influence of intake pressure distortion of the turboshaft engine on aerodynamic stability exists in China, and the calculation model temporarily belongs to the technical blank.
Disclosure of Invention
The invention provides a method, a system and equipment for calculating the aerodynamic stability of a turboshaft engine and a computer readable storage medium, which are used for solving the technical problem that the influence of the intake pressure distortion of the turboshaft engine on the aerodynamic stability cannot be calculated at present.
According to one aspect of the present invention, there is provided a method of calculating aerodynamic stability of a turboshaft engine for evaluating an effect of intake pressure distortion on aerodynamic stability of the turboshaft engine, comprising the steps of:
step S1: constructing a calculation model;
step S2: establishing a calculation area and dividing a control body;
step S3: assuming that the pneumatic parameters within each control volume are spatially uniform, a system of control equations is established to describe the flow within the control volume:
Figure BDA0003022134390000021
wherein p represents a static pressure, T represents a static temperature,
Figure BDA0003022134390000031
denotes the mass flow, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、 r4For the source term, V denotes the control volume, VxIndicating axial velocity, vθRepresents circumferential velocity, v represents total velocity, and t represents time;
step S4: solving the control equation set for each control body to obtain pneumatic parameters at each control body;
step S5: and predicting whether the engine can generate pneumatic instability under the given intake distortion condition based on the pneumatic parameters at each control body and combined with instability criterion.
Further, the step S4 includes the following steps:
step S41: calculating the total pressure ratio, the static pressure ratio and the total temperature ratio of the gas after flowing through the unit by using the unit characteristics;
step S42: calculating the total pressure and the total temperature at each axial section according to the total inlet pressure, the total temperature and the part characteristics of the engine at the initial moment;
step S43: calculating an axial force source term, a circumferential force source term, a power source term and a heat source term according to the total pressure ratio, the total temperature ratio and the static pressure ratio;
step S44: according to the dimensionless dense flow value and the flow continuous equation of the inlet of the engine, calculating the dimensionless dense flow value and the flow of each section under the condition of not considering the influence of bleed air;
step S45: calculating the dimensionless dense flow of the throat part of the tail nozzle according to the boundary condition of the outlet of the engine, and recalculating the area of the throat part of the tail nozzle according to a flow formula;
step S46: assuming that airflow at each axial section in the engine in an initial state flows in a subsonic speed, and calculating the axial speed, the static temperature and the static pressure of each section;
step S47: and calculating the circumferential speed of each section according to the given initial state axial section airflow angle.
Further, the step S43 specifically calculates an axial force source term, a circumferential force source term, a work source term and a heat source term based on the following formulas:
Figure BDA0003022134390000041
Figure BDA0003022134390000042
Figure BDA0003022134390000043
Figure BDA0003022134390000044
wherein, UmThe velocity of the drag at the position of the pitch diameter, beta the relative draft angle, CPWhich represents the specific heat at the same pressure,
Figure BDA0003022134390000045
which is indicative of the air flow rate,
Figure BDA0003022134390000046
indicating fuel flow, HuIndicating low calorific value of fuel, etabIndicating the combustion efficiency.
Further, three boundary conditions are given in the step S4, which are respectively
Inlet total pressure boundary conditions: the steady-state pressure distortion is described by giving total pressure values at different circumferential positions of an inlet section of the engine, and a steady-state pressure distortion index is introduced to express the intensity of the steady-state pressure distortion, wherein the expression is as follows:
Figure BDA0003022134390000047
wherein the content of the first and second substances,
Figure BDA0003022134390000048
which represents the steady-state pressure distortion index,
Figure BDA0003022134390000049
the mean total pressure in the circumferential direction over the entire cross-section is indicated,
Figure BDA00030221343900000410
representing the mean total circumferential pressure of the distortion zone;
inlet airflow direction: giving an axial air inlet boundary condition, namely giving the circumferential speed of inlet airflow to be zero;
exit boundary conditions: the exit boundary is located at the jet nozzle throat section, given the throttling characteristics of the jet nozzle, the formula is as follows:
Figure BDA0003022134390000051
wherein q (lambda) represents a dimensionless dense flow value of the throat section of the jet nozzle,
Figure BDA0003022134390000052
shows the total pressure of the throat part of the tail nozzle
Figure BDA0003022134390000053
Counter pressure p with the outletbThe ratio of (a) to (b).
Further, in step S3, a rotor dynamics equation is also established:
Figure BDA0003022134390000054
wherein the content of the first and second substances,
Figure BDA0003022134390000055
LTrepresenting turbine work, LCWhich is indicative of the compressor's function,
Figure BDA0003022134390000056
which is indicative of the mass flow of the turbine,
Figure BDA0003022134390000057
denotes mass flow of compressor, C'PAnd CPRespectively representing the specific volumetric heat capacity of the air flow passing through the turbine and the compressor,
Figure BDA0003022134390000058
and
Figure BDA0003022134390000059
respectively represent the total temperature ratio obtained after the air flow passes through the turbine and the compressor,
Figure BDA00030221343900000510
and
Figure BDA00030221343900000511
the total inlet temperatures of the turbine and the compressor, respectively.
Further, the step S3 further includes the following steps:
the following control law is adopted to control the oil supply: m isf/p3＝const；
The differential equation of the oil supply regulator corresponding to the control law is as follows:
Figure BDA00030221343900000512
wherein m isfThe flow rate of the fuel is indicated,
Figure BDA00030221343900000513
and (4) representing the relative pressure behind the high-pressure compressor, wherein tau is the time constant of the control system.
Further, the step S2 is specifically:
establishing a calculation region under a cylindrical coordinate system, wherein the coordinate system comprises three coordinate directions: x is axial direction, the airflow flowing direction is positive, theta is circumferential direction and is consistent with the rotating direction of the rotor, r is radial direction, and airflow parameters are assumed to be uniformly distributed along the radial direction, a calculation region comprises the whole flowing region from an engine inlet to the throat part of a tail nozzle, the calculation region is divided into a series of sequentially arranged axial calculation stations by using a plane perpendicular to the axis of the engine under a cylindrical coordinate system, then the calculation region is divided into a plurality of equidistant circumferential calculation stations along the circumferential direction, and then the inner diameter boundary and the outer diameter boundary of an airflow channel are combined to form the closed outer surface of the control body.
In addition, the invention also provides a system for calculating the aerodynamic stability of the turboshaft engine, which comprises the following components:
a model construction unit for constructing a calculation model;
the model dividing unit is used for establishing a calculation area and dividing a control body;
the system of equations establishing unit is used for establishing a system of control equations to describe the flow in the control bodies on the premise that the pneumatic parameters in each control body are uniform in space, and the system of equations is specifically as follows:
Figure BDA0003022134390000061
wherein p represents a static pressure, T represents a static temperature,
Figure BDA0003022134390000062
denotes the mass flow, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、 r4For the source term, V denotes the control volume, VxIndicating axial velocity, vθRepresents circumferential velocity, v represents total velocity, and t represents time;
the calculation unit is used for solving the control equation set aiming at each control body to obtain the pneumatic parameters at each control body;
and the prediction unit is used for predicting whether the engine can generate pneumatic instability under the given intake distortion condition based on the pneumatic parameters at each control body and combined with instability criterion.
In addition, the present invention also provides an apparatus comprising a processor and a memory, wherein the memory stores a computer program, and the processor is used for executing the steps of the method by calling the computer program stored in the memory.
The invention also provides a computer-readable storage medium for storing a computer program for calculating the aerodynamic stability of a turbo machine, wherein the computer program performs the steps of the method described above when the computer program runs on a computer.
The invention has the following effects:
the method for calculating the aerodynamic stability of the turboshaft engine comprises the steps of dividing a calculation model into a plurality of control bodies, calculating a control equation set aiming at each control body to obtain the aerodynamic parameters of each control body, and predicting whether the engine can generate aerodynamic instability under the given intake distortion condition by combining instability criteria, so that the influence of intake pressure distortion on the aerodynamic stability of the turboshaft engine can be accurately evaluated. The invention considers the characteristic that the flow direction of the combined compressor is changed from axial direction to radial direction and then to axial direction, so the adopted calculation model only focuses on the parameters of the inlet and outlet sections of the compressor, and the whole combined compressor can be modularized. In addition, in consideration of the fact that the existence of the power turbine can cause the expansion process of the turbine outlet to change, so that the flow field, the thrust and the oil consumption rate of the outlet of the tail nozzle and the working point position of the turbine can be influenced, therefore, in the calculation model and the calculation method of the turboshaft engine, the power turbine is also used as one calculation unit to participate in the whole iterative calculation process, and the accuracy of the prediction result is improved. And the static pressure, the static temperature, the mass flow and the circumferential speed are used as basic variables to establish the equation set so as to describe the flow in each control body, on one hand, all variables can be obtained through conversion of the equation set formula based on the four basic variables, on the other hand, the four basic variables are convenient to obtain through a test means, and the data obtaining process is simplified.
In addition to the objects, features and advantages described above, other objects, features and advantages of the present invention are also provided. The present invention will be described in further detail below with reference to the drawings.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the invention and, together with the description, serve to explain the invention and are not intended to limit the invention. In the drawings:
FIG. 1 is a flow chart illustrating a method for calculating aerodynamic stability of a turboshaft engine according to a preferred embodiment of the present invention.
FIG. 2 is a computational model of a turboshaft engine constructed in a preferred embodiment of the present invention.
Fig. 3 is a sub-flowchart of step S4 in fig. 1.
FIG. 4 is a graphical representation of the distribution of total pressure values at different axial locations for a given engine inlet cross-section in a preferred embodiment of the present invention.
Fig. 5 is a schematic view of the flow characteristics of an axial-centrifugal compressor in accordance with a preferred embodiment of the present invention.
Fig. 6 is a schematic diagram of the efficiency characteristics of an axial-centrifugal compressor in accordance with a preferred embodiment of the present invention.
Detailed Description
The embodiments of the invention will be described in detail below with reference to the drawings, but the invention can be implemented in many different ways, which are defined and covered below.
As shown in FIG. 1, the preferred embodiment of the present invention provides a method for calculating the aerodynamic stability of a turboshaft engine, for evaluating the influence of intake pressure distortion on the aerodynamic stability of the turboshaft engine, comprising the steps of:
step S1: constructing a calculation model;
step S2: establishing a calculation area and dividing a control body;
step S3: assuming that the pneumatic parameters within each control volume are spatially uniform, a system of control equations is established to describe the flow within the control volume:
Figure BDA0003022134390000091
wherein p represents a static pressure, T represents a static temperature,
Figure BDA0003022134390000092
denotes the mass flow, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、 r4For the source term, V denotes the control volume, VxIndicating axial velocity, vθRepresenting circumferential velocity, v representing velocity, t representing time;
step S4: solving the control equation set for each control body to obtain pneumatic parameters at each control body;
step S5: and predicting whether the engine can generate pneumatic instability under the given intake distortion condition based on the pneumatic parameters at each control body and combined with instability criterion.
It can be understood that, in the method for calculating the aerodynamic stability of the turboshaft engine according to the embodiment, the calculation model is divided into a plurality of control bodies, then the control equation set is calculated for each control body to obtain the aerodynamic parameter at each control body, and then the instability criterion is combined to predict whether the aerodynamic instability of the engine occurs under the given intake distortion condition, so that the influence of the intake pressure distortion on the aerodynamic stability of the turboshaft engine can be accurately evaluated. The invention considers the characteristic that the flow direction of the combined compressor is changed from axial direction to radial direction and then to axial direction, so the adopted calculation model only focuses on the parameters of the inlet and outlet sections of the compressor, and the whole combined compressor can be modularized. In addition, in consideration of the fact that the existence of the power turbine can cause the expansion process of the turbine outlet to change, so that the flow field, the thrust and the oil consumption rate of the outlet of the tail nozzle and the working point position of the turbine can be influenced, therefore, in the calculation model and the calculation method of the turboshaft engine, the power turbine is also used as one calculation unit to participate in the whole iterative calculation process, and the accuracy of the prediction result is improved. And the static pressure, the static temperature, the mass flow and the circumferential speed are used as basic variables to establish the equation set to describe the flow in each control body, on one hand, all variables can be obtained through formula conversion of the equation set on the basis of the four basic variables, on the other hand, the four basic variables are convenient to obtain through a test means, and the data obtaining process is simplified.
A turbo-shaft engine is typically characterized by a combined axial-centrifugal compressor, in addition to a power turbine. In the step S1, a calculation model of the axial-flow centrifugal turboshaft engine is constructed, which is a single-shaft structure, the combined compressor and the gas turbine are on the same shaft, the power turbine is on the other shaft, the combined compressor and the gas turbine are in power balance, the power turbine is in power balance with a load connected behind, and finally, the fixed working point of the engine is realized, and the rotating speed is kept unchanged. Since the rear load is driven by the power turbine, the output power of the power turbine can be adjusted by adjusting the pressure drop ratio and efficiency of the power turbine. Specifically, as shown in fig. 2, the calculation model includes an air inlet pipeline, a combined compressor, a combustion chamber, an axial flow gas turbine, a power turbine, and a tail nozzle, where reference numeral 1 denotes the air inlet pipeline, reference numeral 2 denotes the axial flow + centrifugal combined compressor, reference numeral 3 denotes the combustion chamber, reference numeral 4 denotes the axial flow gas turbine, reference numeral 5 denotes the power turbine, and reference numeral 6 denotes the tail nozzle. Because the parameters of the inlet and outlet sections of the compressor are only concerned in the subsequently adopted control equation set, the whole combined compressor can be modularized without being divided into an axial flow compressor and a centrifugal compressor for separate calculation, and the calculation amount is reduced.
Because the flow area in the aircraft engine has the geometrical characteristic of axial symmetry, the method provided by the invention selects to establish the calculation area and divide the calculation grid under the cylindrical coordinate system. The step S2 specifically includes:
establishing a calculation region under a cylindrical coordinate system, wherein the coordinate system comprises three coordinate directions: x is axial direction, the airflow flowing direction is positive, theta is circumferential direction and is consistent with the rotating direction of the rotor, r is radial direction, and airflow parameters are assumed to be uniformly distributed along the radial direction, a calculation region comprises the whole flowing region from an engine inlet to the throat part of a tail nozzle, the calculation region is divided into a series of sequentially arranged axial calculation stations by using a plane perpendicular to the axis of the engine under a cylindrical coordinate system, then the calculation region is divided into a plurality of equidistant circumferential calculation stations along the circumferential direction, and then the inner diameter boundary and the outer diameter boundary of an airflow channel are combined to form the closed outer surface of the control body. That is, the control body is enclosed by the axial calculation station, the circumferential calculation station, and the inner diameter boundary and the outer diameter boundary of the airflow passage.
In step S3, a control equation system needs to be established for each control body in the calculation region, and the flow inside the control body is described by using the following two-dimensional euler equation with source terms in the form of integral:
the continuous equation:
Figure RE-GDA0003196337760000111
axial momentum equation:
Figure RE-GDA0003196337760000112
circumferential momentum equation:
Figure RE-GDA0003196337760000113
energy equation:
Figure RE-GDA0003196337760000114
wherein g represents the bleed air or gas injection quantity of the compressor or other components, g takes a negative sign when the bleed air is specified, and takes a positive sign when the gas is injected, the lower subscript g represents the bleed air or gas injection parameters, and the subscript theta represents the circumferential componentThe subscript x denotes the axial component, ρ denotes the density,
Figure BDA0003022134390000115
representing the velocity vector, V representing the control volume,
Figure BDA0003022134390000116
representing the area vector, p the pressure, F the volumetric force,
Figure BDA0003022134390000117
the amount of heat is represented by the amount of heat,
Figure BDA0003022134390000118
represents mechanical work, h*Representing total enthalpy, e internal energy, t time.
However, considering that parameters such as flow, pressure and temperature are commonly used in practical engineering to represent the working state of the engine, the invention adopts static pressure p, static temperature T and mass flow
Figure BDA0003022134390000119
And a circumferential velocity VθThese four variables serve as basic variables, and the equal-sign right ends of the four equations in equation set 1 are respectively labeled as r1, r2, r3, r4, while assuming that the pneumatic parameters in the control body are spatially uniform, the above equation set 1 is converted into:
Figure BDA0003022134390000121
wherein p represents a static pressure, T represents a static temperature,
Figure BDA0003022134390000122
representing mass flow rate, VθDenotes the circumferential velocity, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、r4Notation representing the right-hand source term of the above equation set (1), V represents the control volumeVolume, vxRepresenting axial velocity, v representing total velocity, and t representing time.
The above equation set (2) is the control equation set of the calculation model of the present invention.
For simplicity, vectors are introduced
Figure BDA0003022134390000123
Then equation set (2) can be simplified to:
Figure BDA0003022134390000124
in step S3, the above equation set is established using static pressure, static temperature, mass flow and circumferential velocity as basic variables to describe the flow inside each control body, on one hand, all variables can be obtained through conversion of the above equation set formula based on the four basic variables, on the other hand, the four basic variables are conveniently obtained through testing, and the data obtaining process is simplified.
The difficulty in solving equation set (2) for each control volume cell in the calculation region is solving for each source term in the right-hand terms of the equations. The solution method commonly used today is to determine the source term, such as F of a compressor unit, by the CFD methodxAnd FθWhich should be the interaction with the compressor blades as the air flows around them. However, the calculation model of the present invention does not directly calculate these source terms, but rather uses the cell characteristics to calculate the total pressure ratio obtained after flowing through the cell
Figure BDA0003022134390000125
Static pressure ratio pikAnd total temperature ratio theta*Where k denotes the kth computing unit, i.e. the kth control volume, and then the source terms are computed according to different methods. The method does not pay much attention to the flow characteristics in the compressor, but only needs to research the overall characteristics of each stage or all stages of the multistage axial flow compressor, simplifies the acquisition of input data, and can acquire overall performance or test results by numerical calculationAnd the test method is simpler. Calculating the initial conditions, namely the initial conditions, in the model, the initial conditions are steady-state working points of the engine under the uniform air inlet condition, the initial conditions are directly assigned according to the total inlet pressure, the total temperature, the dimensionless dense flow value and the back pressure of the outlet of the tail nozzle at the initial moment, the middle calculation sections are directly assigned by an interpolation method, and the finally established initial conditions are that each calculation station is assigned with the total pressure, the total temperature, the dimensionless dense flow and the static pressure, and the pneumatic parameters, such as the total pressure, the static pressure, the total temperature, the static temperature, the speed and the density, at each axial section of the engine (namely the interface of the axial calculation station divided in the front) are calculated. Specifically, as shown in fig. 3, the step S4 includes the following steps:
step S41: calculating the total pressure ratio, the static pressure ratio and the total temperature ratio of the gas after flowing through the unit by using the unit characteristics;
step S42: calculating the total pressure and the total temperature at each axial section according to the total inlet pressure, the total temperature and the part characteristics of the engine at the initial moment;
step S43: calculating an axial force source term, a circumferential force source term, a power source term and a heat source term according to the total pressure ratio, the total temperature ratio and the static pressure ratio;
step S44: according to the dimensionless dense flow value and the flow continuous equation of the inlet of the engine, calculating the dimensionless dense flow value and the flow of each section under the condition of not considering the influence of bleed air;
step S45: calculating the dimensionless dense flow of the throat part of the tail nozzle according to the boundary condition of the outlet of the engine, and recalculating the area of the throat part of the tail nozzle according to a flow formula;
step S46: assuming that airflow at each axial section in the engine in an initial state flows in a subsonic speed, and calculating the axial speed, the static temperature and the static pressure of each section;
step S47: and calculating the circumferential speed of each section according to the given initial state axial section airflow angle.
In the step S41, the total pressure ratio of the gas after flowing through the cell is calculated by using the cell characteristics
Figure BDA0003022134390000131
Static pressure ratio pi k and total temperature ratio theta*The method specifically comprises the following steps: according to the total pressure ratio and flow characteristic curve, the efficiency and the flow characteristic curve, the flow, the total pressure ratio and the efficiency of the corresponding working point can be directly obtained. According to efficiency
Figure BDA0003022134390000141
And total pressure ratio
Figure BDA0003022134390000142
The total temperature ratio theta can be obtained*. According to the total pressure ratio
Figure BDA0003022134390000143
The static pressure ratio pi can be obtained by summing the velocity coefficient lambdak. The velocity coefficient lambda is calculated according to a flow calculation formula. K is the specific heat ratio, m is the flow rate, q (λ) is the flow rate function, and K is a gas constant, generally equal to 0.04042.
Figure BDA0003022134390000144
Figure BDA0003022134390000145
Figure BDA0003022134390000146
Figure BDA0003022134390000147
In step S42, the total pressure and the total temperature at each axial cross section are calculated according to the calculated total pressure and total temperature of the inlet at the initial time, the total pressure ratio-flow characteristic, and the efficiency-flow characteristic, for example, the compressor outlet parameters are: total pressure
Figure BDA0003022134390000148
Total temperature
Figure BDA0003022134390000149
Static pressure P2＝P1·πk， P1 *Denotes the total inlet pressure, T1 *Denotes the total inlet temperature, P1The inlet static pressure is indicated.
In step S43, the axial force source term, the circumferential force source term, the work source term, and the heat source term are calculated specifically based on the following formulas:
Figure BDA00030221343900001410
Figure BDA0003022134390000151
Figure BDA0003022134390000152
Figure BDA0003022134390000153
wherein the content of the first and second substances,
Figure BDA0003022134390000154
representing the mean value of the axial velocity, theta*Denotes the total temperature ratio,. pi*Representing total pressure ratio,. pi.representing static pressure ratio, k representing kth control body, FxThe axial force is represented by the axial force,
Figure BDA0003022134390000155
denotes the mass flow, subscript m denotes mass, P1Denotes the grid inlet static pressure, A2、A1Respectively representing the cross-sectional areas of the grid outlet and inlet, FθDenotes the circumferential force, CPDenotes the isobaric specific heat, T1 *Indicates the total temperature of grid inlet, UmThe velocity of the drag is indicated at the position of the pitch diameter, and Z is the axisIn the direction of the length of the steel wire,
Figure BDA0003022134390000156
showing the cotangent of the average value of the absolute flow angle of the stator vane discharge port in the compressor unit,
Figure BDA0003022134390000157
denotes the cotangent of the mean value of the relative flow angles of the rotor blade discharge openings, beta denotes the relative flow angle,
Figure BDA0003022134390000158
the mechanical work is represented by the mechanical work,
Figure BDA0003022134390000159
it is meant that the amount of heat,
Figure BDA00030221343900001510
which is indicative of the air flow rate,
Figure BDA00030221343900001511
indicating fuel flow, CP2Representing the isobaric specific heat of the outlet gas,
Figure BDA00030221343900001512
indicating total temperature of compressor outlet, CP1Denotes the isobaric specific heat, H, of the inlet gasuIndicating low calorific value of fuel, etabIndicating combustion efficiency, CtExpressing specific heat capacity of fuel vaporization, TfIndicating the fuel temperature.
It can be understood that step S44 specifically includes:
(1) the air flow in the air inlet channel and the air compressor is kept unchanged;
(2) the gas flow at the outlet of the combustion chamber is equal to the air flow at the inlet of the combustion chamber plus the fuel flow;
(3) the gas flow of the turbine and the tail nozzle are equal.
It can be understood that step S45 specifically includes:
(1) nozzle velocity characteristics refer to total throat pressure in terms of engine nozzle
Figure BDA0003022134390000161
Counter pressure P with the outletbCalculating the dimensionless dense flow q (lambda) of the nozzle throat2) The formula is as follows:
Figure BDA0003022134390000162
the formula is used for approximately calculating the dimensionless dense flow of the throat part of the spray pipe according to the ratio of the total pressure to the back pressure of the inlet of the spray pipe. If it is assumed that the throat static pressure is Pb(the formation of subsonic convergent nozzle)
When in use
Figure BDA0003022134390000163
The dimensionless dense flow calculated from the total static pressure ratio is equal to 1. With following
Figure BDA0003022134390000164
Further increasing, the dimensionless dense flow calculated according to the total static pressure ratio is gradually reduced, and the speed characteristic of the spray pipe can be better simulated.
(2) Determining the area of the throat of the exhaust nozzle
Throat area of steady state operating point
Figure BDA0003022134390000165
(i.e., convergent nozzle orifice area), the calculation is as follows:
Figure BDA0003022134390000166
in the formula: the constant of the K flow function is,
Figure BDA0003022134390000167
G2mass flow at throat, kg/s;
Figure BDA0003022134390000168
total pressure in the throat, Pa；
T1 *Total throat temperature, K;
q(λ2) The throat is dimensionless dense flow.
It can be understood that step S46 specifically includes:
(1) solving a speed coefficient lambda according to a flow function q (lambda):
Figure BDA0003022134390000169
(2) according to the total pressure P*And static pressure P is obtained by the relation between:
Figure BDA0003022134390000171
(3) according to the total temperature T*And the static temperature T, obtaining the static pressure T according to the relation formula:
Figure BDA0003022134390000172
(4) when the velocity is in the axial direction, the axial velocity is obtained from the relationship between the velocity coefficient λ and the axial velocity Vz, where R is a gas constant and k is a specific heat ratio.
Figure BDA0003022134390000173
Three boundary conditions are also given during calculation, namely an inlet total pressure boundary condition, an inlet airflow direction and an outlet boundary condition. The calculation model of the invention has the remarkable characteristic that the comprehensive pressure distortion index can be used for describing the total pressure distortion of the inlet, so that the total pressure condition of the inlet comprises two parts, one part is used for describing steady-state (ordered) distortion, and the other part is used for describing random dynamic distortion:
Figure BDA0003022134390000174
while steady state (ordered) circumferential pressure distortion
Figure BDA0003022134390000175
The method also comprises two parts of steady-state pressure distortion and unsteady pressure distortion:
Figure BDA0003022134390000176
the first term p on the right of the middle symbol in the formula (5)*(θ) is the steady state pressure distortion. In the present calculation model, the steady state pressure distortion is described by the total pressure values at different circumferential positions of a given engine inlet cross section, which is distributed as shown in fig. 4. In the context of figure 4 of the drawings,
Figure BDA0003022134390000177
the mean total pressure in the circumferential direction over the entire cross-section is indicated,
Figure BDA0003022134390000178
represents the circumferential average total pressure of the low pressure zone (i.e., the distortion zone) and Δ θ represents the circumferential extent of the distortion zone. The second term on the right of the equal sign of formula (5) is unsteady pressure distortion and can be used for describing shock waves, wherein dp*The rate at which the total pressure varies with time is/dt and can be dependent on the circumferential position (two-dimensional) or independent of the circumferential position (one-dimensional), t0Indicating the starting moment at which the total pressure starts to change.
The invention introduces the steady state pressure distortion index
Figure BDA0003022134390000181
To express the intensity of the steady state pressure distortion, the expression is as follows:
Figure BDA0003022134390000182
in addition, given the axial inlet boundary condition, i.e. given a circumferential component velocity of the inlet flow of zero:
vθ(θ,t)＝0 (7)
exit boundary conditions: the exit boundary is located at the jet nozzle throat section, given the throttling characteristics of the jet nozzle, the formula is as follows:
Figure BDA0003022134390000183
wherein q (lambda) represents a dimensionless dense flow value of the throat section of the jet nozzle,
Figure BDA0003022134390000184
shows the total pressure of the throat part of the tail nozzle
Figure BDA0003022134390000185
Ratio to outlet back pressure pb.
Preferably, in order to take account of the influence of the fuel regulation system, the control equation set is supplemented with a rotor dynamics equation, i.e. the rotor dynamics equation is also established in step S3:
Figure BDA0003022134390000186
wherein M isTAnd MCRespectively representing the torque acting on the turbine and compressor rotors, omega being the angular frequency of rotation of the rotors and I being the moment of inertia of the rotors.
And the torque difference can be expressed by the difference between the turbine work and the compressor work, and the rotation angle frequency can be expressed by the rotation speed, the above equation (9) can be rewritten as
Figure BDA0003022134390000191
Wherein the content of the first and second substances,
Figure BDA0003022134390000192
LTrepresenting turbine work, LCWhich is indicative of the compressor's function,
Figure BDA00030221343900001913
which is indicative of the mass flow of the turbine,
Figure BDA0003022134390000193
indicating the mass flow of the compressor, C'PAnd CPRespectively representing the specific volumetric heat capacity of the gas streams passing through the turbine and the compressor,
Figure BDA0003022134390000194
and
Figure BDA0003022134390000195
respectively representing the total temperature ratio obtained after the air flow passes through the turbine and the compressor,
Figure BDA0003022134390000196
and
Figure BDA0003022134390000197
the total inlet temperatures of the turbine and the compressor, respectively.
In addition, step S3 provides three control laws to control the oil supply amount, which are:
1) a power turbine rotor: n is1＝const；
2) A gas turbine rotor: n is2＝const；
The differential equation of the oil supply regulator corresponding to the two control laws is as follows:
Figure BDA0003022134390000198
wherein the content of the first and second substances,
Figure BDA0003022134390000199
in order to be the relative fuel flow rates,
Figure BDA00030221343900001910
for the relative rotational speed of the rotors, i-1 or 2 corresponds to the power turbine rotor or the gas, respectivelyThe turbine rotor, a and b are constants, describing the inertia of the control system.
3)mf/p3＝const，p3Is the pressure behind the high-pressure compressor,
the differential equation of the oil supply regulator corresponding to the control law is as follows:
Figure BDA00030221343900001911
wherein m isfThe flow rate of the fuel is indicated,
Figure BDA00030221343900001912
and (4) representing the relative pressure behind the high-pressure compressor, wherein tau is the time constant of the control system.
The step S4 specifically includes: and solving the equation (3) by adopting a four-order Runge-Kutta explicit method. Suppose that
Figure BDA0003022134390000201
Is the solution of equation (3) at the ith time step, then the solution at the (i + 1) th time step
Figure BDA0003022134390000202
Can be determined by the following formula:
Figure BDA0003022134390000203
and, instead,
Figure BDA0003022134390000204
wherein h represents an integration step length, which is generally 0.0002-0.0005 s, and the stability of the calculation process and the accuracy of the calculation result can be ensured by using the integration step length.
The method is adopted to solve the flow in the calculation area to obtain the pneumatic parameters of each control body in the engine, and then reliable and accurate instability criterion is combined to predict whether the engine can generate pneumatic instability under the given intake distortion condition. In the literature published in China at present, the common method is to use the negative axial speed of a certain section of an engine as a criterion for judging the instability of the engine.
In practical application, before calculation, the equal rotation speed characteristics of the combined compressor need to be given, including 6 equal rotation speed lines, and a flow function, a total pressure ratio and efficiency of 15 working points are given on each equal rotation speed line, as shown in fig. 5 and 6.
Table 1 shows the flow, total pressure ratio and efficiency of the operating point and stable boundary point when the relative reduced rotation speed is equal to 1.0, 0.95 and 0.9, and the stability margin SM of three equal rotation speeds is obtained by calculation according to these parameters, and the calculation formula is as follows:
Figure BDA0003022134390000211
wherein, pi*The total pressure ratio is expressed as,
Figure BDA0003022134390000212
indicating the reduced flow, the subscript s indicating the stable boundary point and the subscript w indicating the operating point. The stability margins at relative reduced rotational speeds equal to 1.0, 0.95 and 0.9 are 20.634%, 12.469% and 11.34%, respectively. High pressure gas is extracted at the combined compressor outlet for cooling the gas turbine nozzle and rotor, the extracted cooling gas accounting for 13% of the total air flow, of which 7% is used for cooling the gas turbine nozzle and 6% is used for cooling the gas turbine rotor.
When the model engine intake pressure distortion calculation analysis is carried out, the engine is divided into 2 circumferential sectors, wherein 1 sector is a low-pressure area, and the other 1 sector is a high-pressure area, namely the low-pressure area range is 180 degrees. Under inlet distortion conditions, the drop pressure ratio of the power turbine decreases, and at 100%, 95%, and 90% design speed, the drop pressure ratio of the power turbine is equal to 2.535, 2.389, and 2.219, respectively. The critical pressure distortion index and the pressure distortion sensitivity coefficient under different rotating speeds are given in table 1, along with the reduction of the rotating speed, the original available stability margin of the engine is gradually reduced, the critical comprehensive pressure distortion index is also gradually reduced, and the pressure distortion sensitivity coefficient is gradually increased, which indicates that the pressure distortion resistance of the engine is gradually weakened. The magnitude of the anti-pressure distortion capability of the engine is related to the steepness of the characteristic line of the compressor, and as can be seen from the combined compressor characteristic diagram in fig. 5, as the rotating speed is reduced, the absolute value of the slope of the equal rotating speed line at the position of the common working point is gradually reduced, the steepness is reduced, and the characteristic line is more and more gentle, so that the anti-pressure distortion capability of the engine is reduced. Therefore, in order to improve the pressure distortion resistance of the engine, it is necessary to increase the slope of the common operating point on the equal rotation speed line as much as possible.
TABLE 1 Critical integrated pressure distortion index and pressure distortion sensitivity coefficient under different rotation speeds
Figure BDA0003022134390000213
In addition, the present invention also provides a system for calculating the aerodynamic stability of a turboshaft engine, preferably using the method for calculating the aerodynamic stability of a turboshaft engine as described above, the system comprising:
a model construction unit for constructing a calculation model;
the model dividing unit is used for establishing a calculation area and dividing a control body;
the system of equations establishing unit is used for establishing a system of control equations to describe the flow in the control bodies on the premise that the pneumatic parameters in each control body are uniform in space, and the system of equations is specifically as follows:
Figure BDA0003022134390000221
wherein p represents a static pressure, T represents a static temperature,
Figure BDA0003022134390000222
denotes the mass flow, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、 r4For the source term, V denotes the control volume, VxIndicating axial velocity, vθRepresenting circumferential velocity, v representing velocity, t representing time;
the calculation unit is used for solving the control equation set aiming at each control body to obtain the pneumatic parameters at each control body;
and the prediction unit is used for predicting whether the engine can generate pneumatic instability under the given intake distortion condition based on the pneumatic parameters at each control body and combined with instability criterion.
It can be understood that the working principle and process of each unit in the system of this embodiment correspond to each step in the above method embodiment, and therefore are not described herein again.
The system for calculating the aerodynamic stability of the turboshaft engine in the embodiment can accurately evaluate the influence of intake pressure distortion on the aerodynamic stability of the turboshaft engine by dividing a calculation model into a plurality of control bodies, then calculating a control equation set aiming at each control body to obtain the aerodynamic parameters of each control body, and then predicting whether the aerodynamic instability of the engine can occur under the given intake distortion condition by combining instability criteria. The invention considers the characteristic that the flow direction of the combined compressor is changed from axial direction to radial direction and then to axial direction, so the adopted calculation model only focuses on the parameters of the inlet and outlet sections of the compressor, and the whole combined compressor can be modularized. In addition, the existence of the power turbine can cause the expansion process of the turbine outlet to change, so that the outlet flow field, the thrust and the oil consumption rate of the tail nozzle and the working point position of the turbine can be influenced, therefore, in the calculation model and the calculation method of the turboshaft engine, the power turbine is also used as one calculation unit to participate in the whole calculation iteration process, and the accuracy of the prediction result is improved. And static pressure, static temperature, mass flow and circumferential speed are used as basic variables to establish the equation set to describe the flow in each control body, on one hand, all variables can be obtained through formula conversion of the equation set on the basis of the four basic variables, on the other hand, the four basic variables are convenient to obtain through a test means, and the data obtaining process is simplified.
In addition, the present invention also provides an apparatus comprising a processor and a memory, wherein the memory stores a computer program, and the processor is used for executing the steps of the method by calling the computer program stored in the memory.
In addition, the present invention also provides a computer readable storage medium for storing a computer program for calculating the aerodynamic stability of a turboshaft engine, which computer program, when running on a computer, performs the steps of the method as described above.
The general form of computer readable media includes: floppy disk (floppy disk), flexible disk (flexible disk), hard disk, magnetic tape, any other magnetic medium, CD-ROM, any other optical medium, punch cards (punch cards), paper tape (paper tape), any other physical medium with patterns of holes, Random Access Memory (RAM), Programmable Read Only Memory (PROM), Erasable Programmable Read Only Memory (EPROM), FLASH erasable programmable read only memory (FLASH-EPROM), any other memory chip or cartridge, or any other medium from which a computer can read. The instructions may further be transmitted or received by a transmission medium. The term transmission medium may include any tangible or intangible medium that is operable to store, encode, or carry instructions for execution by the machine, and includes digital or analog communications signals or any intangible medium that facilitates communication of such instructions. Transmission media include coaxial cables, copper wire and fiber optics, including the wires that comprise a bus for transmitting a computer data signal.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, improvement and the like made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims (10)

1. A method of calculating aerodynamic stability of a turboshaft engine for assessing the effect of inlet pressure distortion on aerodynamic stability of the turboshaft engine, comprising the steps of:
step S1: constructing a calculation model;
step S2: establishing a calculation area and dividing a control body;
step S3: assuming that the pneumatic parameters within each control volume are spatially uniform, a system of control equations is established to describe the flow within the control volume:
Figure FDA0003022134380000011
wherein p represents a static pressure, T represents a static temperature,
Figure FDA0003022134380000012
denotes the mass flow, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、r4For the source term, V denotes the control volume, VxIndicating axial velocity, vθRepresenting circumferential velocity, v representing total velocity, t representing time;
step S4: solving the control equation set for each control body to obtain pneumatic parameters at each control body;
step S5: and predicting whether the engine can generate pneumatic instability under the given intake distortion condition based on the pneumatic parameters at each control body and combined with instability criterion.
2. The method of calculating aerodynamic stability of a turboshaft engine according to claim 1,
the step S4 includes the steps of:
step S41: calculating the total pressure ratio, the static pressure ratio and the total temperature ratio of the gas after flowing through the unit by using the unit characteristics;
step S42: calculating the total pressure and the total temperature at each axial section according to the total inlet pressure, the total temperature and the part characteristics of the engine at the initial moment;
step S43: calculating an axial force source term, a circumferential force source term, a power source term and a heat source term according to the total pressure ratio, the total temperature ratio and the static pressure ratio;
step S44: according to the dimensionless dense flow value and the flow continuous equation of the inlet of the engine, calculating the dimensionless dense flow value and the flow of each section under the condition of not considering the influence of bleed air;
step S45: calculating the dimensionless dense flow of the throat part of the tail nozzle according to the boundary condition of the outlet of the engine, and recalculating the area of the throat part of the tail nozzle according to a flow formula;
step S46: assuming that airflow at each axial section in the engine in an initial state flows in a subsonic speed, and calculating the axial speed, the static temperature and the static pressure of each section;
step S47: and calculating the circumferential speed of each section according to the given initial state axial section airflow angle.
3. The method of calculating aerodynamic stability of a turboshaft engine according to claim 2,
step S43 is to calculate an axial force source term, a circumferential force source term, a work source term, and a heat source term based on the following formulas:
Figure FDA0003022134380000021
Figure FDA0003022134380000022
Figure FDA0003022134380000031
Figure FDA0003022134380000032
wherein, UmThe velocity of the drag at the position of the pitch diameter, beta the relative draft angle, CPWhich represents the specific heat at the same pressure,
Figure FDA0003022134380000033
which is indicative of the air flow rate,
Figure FDA0003022134380000034
indicating fuel flow, HuIndicating low calorific value of fuel, etabIndicating the combustion efficiency.
4. The method of calculating aerodynamic stability of a turboshaft engine according to claim 2,
in the step S4, three boundary conditions are given, namely
Inlet total pressure boundary conditions: the steady-state pressure distortion is described by giving total pressure values at different circumferential positions of an inlet section of the engine, and a steady-state pressure distortion index is introduced to express the intensity of the steady-state pressure distortion, wherein the expression is as follows:
Figure FDA0003022134380000035
wherein the content of the first and second substances,
Figure FDA0003022134380000036
which represents the steady-state pressure distortion index,
Figure FDA0003022134380000037
the mean total pressure in the circumferential direction over the entire cross-section is indicated,
Figure FDA0003022134380000038
representing the mean total circumferential pressure of the distortion zone;
inlet airflow direction: giving an axial air inlet boundary condition, namely giving the circumferential component velocity of the inlet airflow to be zero;
exit boundary conditions: the exit boundary is located at the jet nozzle throat section, given the throttle characteristics of the jet nozzle, the formula is as follows:
Figure FDA0003022134380000039
wherein q (lambda) represents a dimensionless dense flow value of the throat section of the jet nozzle,
Figure FDA0003022134380000041
indicating total pressure at the throat of the exhaust nozzle
Figure FDA0003022134380000042
Counter pressure p with the outletbThe ratio of (a) to (b).
5. The method of calculating aerodynamic stability of a turboshaft engine according to claim 1,
in step S3, a rotor dynamics equation is also established:
Figure FDA0003022134380000043
wherein the content of the first and second substances,
Figure FDA0003022134380000044
LTrepresenting turbine work, LCWhich is indicative of the compressor's function,
Figure FDA0003022134380000045
which is indicative of the mass flow of the turbine,
Figure FDA0003022134380000046
denotes mass flow of compressor, C'PAnd CPRespectively representing the specific volumetric heat capacity of the air flow passing through the turbine and the compressor,
Figure FDA0003022134380000047
and
Figure FDA0003022134380000048
respectively representing the total temperature ratio obtained after the air flow passes through the turbine and the compressor,
Figure FDA0003022134380000049
and
Figure FDA00030221343800000410
the total inlet temperatures of the turbine and the compressor, respectively.
6. The method of calculating aerodynamic stability of a turboshaft engine according to claim 5,
the step S3 further includes the following steps:
the following control law is adopted to control the oil supply: m isf/p3＝const；
The differential equation of the oil supply regulator corresponding to the control law is as follows:
Figure FDA00030221343800000411
wherein m isfThe flow rate of the fuel is indicated,
Figure FDA00030221343800000412
and (4) representing the relative pressure behind the high-pressure compressor, wherein tau is the time constant of the control system.
7. The method of calculating aerodynamic stability of a turboshaft engine according to claim 1,
the step S2 specifically includes:
establishing a calculation region under a cylindrical coordinate system, wherein the coordinate system comprises three coordinate directions: x is axial direction, the airflow flowing direction is positive, theta is circumferential direction and is consistent with the rotating direction of the rotor, r is radius direction, and the airflow parameters are assumed to be uniformly distributed along the radial direction, the calculation region comprises the whole flowing region from the inlet of the engine to the throat part of the tail nozzle, the calculation region is divided into a series of sequentially arranged axial calculation stations by using a plane vertical to the axis of the engine under a cylindrical coordinate system, then the calculation region is divided into a plurality of equidistant circumferential calculation stations along the circumferential direction, and then the inner diameter boundary and the outer diameter boundary of the airflow channel are combined to form the closed outer surface of the control body.
8. A system for calculating aerodynamic stability of a turboshaft engine, comprising:
a model construction unit for constructing a calculation model;
the model dividing unit is used for establishing a calculation area and dividing a control body;
the system of equations establishing unit is used for establishing a control system of equations to describe the flow in each control body on the premise that the pneumatic parameters in each control body are uniform in space, and the system of equations is specifically as follows:
Figure FDA0003022134380000051
wherein p represents a static pressure, T represents a static temperature,
Figure FDA0003022134380000052
denotes the mass flow, L denotes the length, R denotes the gas constant, the index g denotes the bleed or injection gas parameter, CvDenotes the specific heat of constant heat capacity, r1、r2、r3、r4For the source term, V denotes the control volume, VxIndicating axial velocity, vθRepresenting circumferential velocity, v representing total velocity, t representing time;
the calculation unit is used for solving the control equation set aiming at each control body to obtain the pneumatic parameters at each control body;
and the prediction unit is used for predicting whether the engine can generate pneumatic instability under the given intake distortion condition based on the pneumatic parameters at each control body and combined with instability criterion.
9. An apparatus comprising a processor and a memory, the memory having stored therein a computer program, the processor being configured to perform the steps of the method of any one of claims 1 to 7 by invoking the computer program stored in the memory.
10. A computer-readable storage medium for storing a computer program for calculating the aerodynamic stability of a turboshaft engine, wherein the computer program, when run on a computer, performs the steps of the method according to any one of claims 1 to 7.
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罗志煌: "涡轴发动机稳定性实验研究", 《推进技术》 *

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CN114036449A (en) * 2021-10-15 2022-02-11 北京航空航天大学 Thermo-acoustic stability prediction method and device
CN116151157A (en) * 2023-04-23 2023-05-23 中国航发四川燃气涡轮研究院 Calculation method for simulating surge hammering wave load of engine
CN116151157B (en) * 2023-04-23 2023-06-30 中国航发四川燃气涡轮研究院 Calculation method for simulating surge hammering wave load of engine
CN117313579A (en) * 2023-10-07 2023-12-29 中国航空发动机研究院 Engine compression part flow field prediction method, device, equipment and storage medium
CN117313579B (en) * 2023-10-07 2024-04-05 中国航空发动机研究院 Engine compression part flow field prediction method, device, equipment and storage medium
CN117252109A (en) * 2023-11-14 2023-12-19 太仓点石航空动力有限公司 Aeroengine stability analysis method and system based on data processing
CN117252109B (en) * 2023-11-14 2024-01-26 太仓点石航空动力有限公司 Aeroengine stability analysis method and system based on data processing

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