Turbine engine matching condition assessment method considering air system swirl transfer

Abstract

The present invention discloses a method for evaluating the matching state of a turbine engine taking into account the swirl transfer of the air system, which is applied to the technical field of evaluation of the operating state of aviation turbine engines. The present invention comprises: acquiring engine data and preprocessing the engine data; establishing a whole-machine transient model that describes the conservation relationship of aerodynamic thermodynamic matching in the engine based on the preprocessed engine data and taking into account the swirl transfer of the engine air system; solving the established whole-machine transient model to obtain engine operating state parameters and realize engine matching state evaluation. The present invention realizes the automatic calculation of the air system inlet / exhaust parameters and the transmission of the air system swirl parameters along the flow direction; the established whole-machine transient model can be used for the evaluation of the aerodynamic thermodynamic matching state of the aviation turbine engine in steady state and transition process, which greatly improves the efficiency of the air system parameter calculation and has high calculation accuracy.

Description

Technical Field

[0001] The present invention relates to the technical field of aviation turbine engine operating state evaluation, and more particularly to a turbine engine matching state evaluation method considering air system swirl transfer. Background Art

[0002] Aircraft turbine engines are rotating machines whose internal flow and heat transfer are influenced by rotation, ultimately impacting engine performance and safety. In particular, within the air system of an aircraft turbine engine, numerous swirl-dominated processes, such as disc cavity flow, exist. Therefore, it is necessary to consider swirl in the design and analysis of aircraft turbine engines, particularly air systems. However, this process exhibits two coupled effects. First, the inlet / exhaust parameters at the interface between the air system and the main flow passage are determined by the operating state of the turbomachinery, while the inlet and outlet boundary conditions of the turbomachinery are typically determined by the matching operating point of the engine's main flow passage components. Second, while the inlet / exhaust parameters define the boundaries of the air system, parameters such as the swirl coefficient are transmitted along the air system. Through effects such as swirl pressure rise and windage temperature rise, they significantly influence the aerodynamic thermodynamic parameters within the air system. This, in turn, influences the matching operating point of the main flow passage components through the air system's inlet / exhaust ratio. Consequently, under various off-design and transient conditions, this coupled process makes it difficult to efficiently and quantitatively calculate the engine's matching operating point and parameters such as the swirl coefficient within the air system. This hinders the design and analysis of aircraft turbine engines and their air systems. For the whole-machine modeling of aviation turbine engines and the calculation of air system parameters, the existing technologies mainly include the following two types: (1) three-dimensional modeling of the impeller machinery combined with air system fluid network modeling and solution with local swirl calculation; (2) full three-dimensional modeling and solution of the whole machine.

[0003] There are at least the following problems with the existing technology:

[0004] (1) It is difficult to automatically match and calculate the inlet / exhaust parameters at the intersection of the air system and the main channel. For the three-dimensional modeling of the impeller machinery combined with the fluid network modeling and solution method of the air system with local swirl calculation, it is necessary to first obtain the aerodynamic thermodynamic boundary conditions of the inlet and outlet of the impeller machinery through the overall performance calculation model of the aviation turbine engine; on this basis, the aerodynamic thermodynamic parameters at the inlet / exhaust position inside the impeller machinery are calculated through the three-dimensional model of the impeller machinery; further, the parameters inside the air system along the flow direction are calculated through the fluid network model. However, there is a coupling effect in the above process, and it is necessary to solve it together or iteratively to obtain the accurate engine matching working point, inlet / exhaust parameters and internal parameters of the air system. When it is necessary to consider multiple non-design points or the transient process response of the engine, the above method is difficult to automatically match and calculate the inlet / exhaust parameters at the intersection of the air system and the main channel.

[0005] (2) It is difficult to achieve swirl parameter transfer between components within the air system. Current air system fluid network modeling methods often only consider swirl effects in areas where swirl effects are significant, such as the disc cavity, and the inlet swirl ratio is generally determined based on engineering experience. This approach not only ignores the swirl parameter transfer relationship between upstream and downstream components within the air system, but also makes it difficult to consider the impact of inertia on swirl parameter changes during engine transitions.

[0006] (3) Low computational efficiency. The three-dimensional modeling of the impeller machinery combined with the modeling and solution of the air system fluid network with local swirl calculations involves models of multiple disciplines and time and space scales, which require iterative solutions. The full three-dimensional modeling and solution of the entire machine requires a huge amount of grid, which consumes a lot of computing resources.

[0007] Therefore, it is an urgent problem for those skilled in the art to propose a turbine engine matching status assessment method that takes into account the swirl transfer of the air system to solve the difficulties existing in the prior art. Summary of the Invention

[0008] In view of this, the present invention provides a turbine engine matching state evaluation method considering the swirl transfer of the air system. The established whole-machine transient model can be used to evaluate the aerodynamic thermodynamic matching state of aviation turbine engines in steady state and transition process.

[0009] In order to achieve the above object, the present invention provides the following technical solutions:

[0010] A method for evaluating a turbine engine matching state considering air system swirl transfer comprises the following steps:

[0011] S1. Obtain engine data and pre-process the engine data;

[0012] S2. Based on the engine data preprocessed in S1, considering the swirl transfer of the engine air system, a whole-machine transient model describing the aerodynamic thermodynamic matching conservation relationship in the engine is established;

[0013] S3. Solve the whole machine transient model established in S2, obtain the engine operating state parameters, and realize the engine matching state evaluation.

[0014] Optionally, the turbine engine model is established in S2 as follows:

[0015] The turbine engine is modeled as a combination of nodes and elements. As a fluid-thermal machine, the turbine engine has node types including compressible fluid nodes, thermal nodes, and mechanical nodes. Among them, the compressible fluid nodes are divided into two types of nodes: main channel and air system, depending on whether the swirl parameter transfer is considered. The elements include main channel elements, inlet / exhaust elements, and air system elements.

[0016] Optionally, the conservation equations for the mass, energy, water vapor component, and oil vapor component of the compressible fluid at the main flow channel node are as follows:

[0017]

[0018] Where V represents the volume of the node, ρ represents the density of the compressible fluid, T represents the temperature of the compressible fluid, c v represents the specific heat ratio of compressible gas at constant volume, f fuel represents the oil and gas mass fraction of the compressible fluid, f water represents the water vapor mass fraction of the compressible fluid, τ represents time; n represents the number of elements connected to the node, Represents the mass flow into or out of a node.

[0019] Optionally, based on the conservation equations of the main channel nodes, the air system nodes also have the following angular momentum conservation relationship:

[0020]

[0021] Where r represents the average rotation radius at the node, V θ represents the angular velocity of the compressible fluid.

[0022] Optionally, the transfer process in the solid domain of the engine is modeled as a thermal node, where the following energy conservation relationship exists:

[0023]

[0024] Among them, c p represents the constant-pressure specific heat of compressible gas, Q i is the node heat flow, which is positive when flowing into the node and negative when flowing out of the node.

[0025] Optionally, the energy conservation relationship between the power generation end and the power consumption end is implemented at the mechanical node as follows:

[0026]

[0027] Where ω is the angular velocity of the rotor, J is the moment of inertia of the rotor, and W i Represents the transferred power.

[0028] Optionally, the governing equations of the air system components represent the flow rate, windage temperature rise, composition, and swirl relationship at the inlet and outlet of the components, respectively:

[0029]

[0030]

[0031] Among them, the subscripts in and out represent the inlet and outlet of the component respectively; represents the mass flow rate of the fluid flowing through the element, T represents the temperature of the compressible fluid, c p represents the constant-pressure specific heat of compressible gas, f fuel represents the oil and gas mass fraction of the compressible fluid, f water represents the water vapor mass fraction of the compressible fluid, r represents the average rotation radius of the element, V θ represents the angular velocity of the compressible fluid; M rotor ω represents the windage torque, Q c Represents the heat exchange between the component and the thermal node; M rotor ,M stator represent the torques exerted by the rotor and stator on the airflow respectively.

[0032] Optionally, first define the enthalpy rise ratio coefficient of the turbomachinery:

[0033]

[0034] Among them, H in ,H out ,H bleed The separate tables represent the total enthalpy of the fluid at the inlet, outlet, and exhaust positions of the impeller machinery. Based on the inlet temperature, inlet pressure, and outlet pressure of the impeller machinery, the total enthalpy at the outlet of the impeller machinery is obtained by interpolation of the characteristic diagram, and then the total enthalpy at the exhaust position of the impeller machinery is obtained based on the enthalpy rise ratio coefficient:

[0035] H bleed =α·(H out -H in )+Hx n

[0036] The total pressure at the impeller machine inlet / exhaust position is calculated through the adiabatic process:

[0037]

[0038] Among them, P t,in ,T t,in ,P t,bleed ,T t,bleed are the total pressure and total temperature at the turbomachinery inlet and introduction / exhaust positions, respectively, and γ represents the polynomial index of this thermodynamic process.

[0039] Optionally, based on the aerodynamic thermodynamic function, the axial velocity coefficient λ at the inlet / exhaust position can be calculated axial :

[0040]

[0041] Among them, q(λaxial ) is the dimensionless density flow of the main channel gas, is the mass flow rate of the gas in the main channel, A is the effective flow area of ​​the gas channel, K is a constant related to the specific heat ratio k of the gas; P t,bleed ,T t,bleed are the total pressure and total temperature at the turbine machine inlet / exhaust positions respectively;

[0042] The circumferential velocity coefficient is obtained by trigonometric function:

[0043] λ θ,vane =λ axial tan(θ)

[0044] Among them, λ θ,vane ,λ axial are the relative circumferential velocity coefficient and axial velocity coefficient respectively, and θ is the angle between the outlet position of the blade structure and the engine axis;

[0045] For the airflow angle behind the guide vane, the swirl coefficient in the stationary coordinate system is λ θ，vane , and for the rotating rotor blades, the coordinate system is transformed to obtain the swirl velocity in the stationary coordinate system:

[0046]

[0047] Among them, λ θ,blade ,λ θ,vane represent the absolute circumferential velocity coefficient and the relative circumferential velocity coefficient respectively; is the rotor rotation speed coefficient, ω is the rotor speed, r is the radius at the inlet / exhaust position, R is the gas constant, k is the specific heat ratio of the gas, T t,bleed are the total temperatures at the turbine inlet / exhaust locations respectively;

[0048] Calculate the swirl velocity at the inlet / exhaust position:

[0049]

[0050] Among them, V θ ,λ θ are the circumferential velocity and circumferential velocity coefficient of the gas, R is the gas constant, k is the specific heat ratio of the gas, T t,bleed are the total temperatures at the turbine inlet / exhaust locations respectively;

[0051] The total velocity coefficient at the inlet / exhaust position is the composite velocity coefficient of the axial and circumferential directions:

[0052]

[0053] Among them, λ bled ,λ θ ,λaxial are the total velocity coefficient, circumferential velocity coefficient and axial velocity coefficient at the inlet / exhaust position respectively;

[0054] The static pressure at the inlet / exhaust position is obtained by the aerodynamic thermodynamic relationship:

[0055] P s,bleed =P t,bleed ·π(λ bleed )

[0056] Among them, P s,bleed ,P t,bleed are the static pressure and total pressure at the inlet / exhaust position, respectively, bleed is the total velocity coefficient at the inlet / exhaust position, π(λ bleed ) is the pressure ratio function.

[0057] As can be seen from the above technical solutions, compared with the prior art, the present invention discloses a method for evaluating the matching status of a turbine engine taking into account the swirl transfer of the air system, which has the following beneficial effects:

[0058] (1) The present invention automatically matches the inlet / exhaust parameters at the intersection of the air system and the main flow channel. Automatic matching of the total temperature, total pressure, static pressure, and swirl velocity at the inlet / exhaust position is achieved through the theoretical method of the impeller machinery. The inlet / exhaust parameters are related to the operating state of the impeller machinery or the matching operating point of the entire machine. Therefore, the global coupling relationship of the entire machine is considered, and a coupled calculation model of the main flow channel and the air system is established. The inlet / exhaust parameters at the intersection of the air system and the main flow channel can be automatically matched under different engine operating conditions and during transition processes.

[0059] (2) Realize the transfer of swirl parameters between various components within the air system. Model the entire engine as a combination of nodes and elements, introduce the angular momentum conservation equation into the control equation, consider the effects of swirl pressure rise and windage temperature rise on the components, and consider the swirl transfer relationship between various components within the air system in the overall matching.

[0060] (3) Improved computational efficiency. The matching calculation is based on the lumped parameter network model of the entire machine; the main flow channel and air system state parameters are solved using a unified form of conservation equations; the inlet / exhaust parameters are based on theoretical formulas, rather than relying on high-dimensional models of the impeller machinery, which reduces the requirements for matching computing resources. BRIEF DESCRIPTION OF THE DRAWINGS

[0061] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are merely embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying any creative work.

[0062] Figure 1 A flow chart of a turbine engine matching status assessment method considering air system swirl transfer provided by the present invention;

[0063] Figure 2 A schematic diagram of the classification of turbine engine model modules of the present invention;

[0064] Figure 3 Schematic diagram of the coupling relationship of component parameters of the present invention;

[0065] Figure 4 It is a schematic diagram of the engine modeling application of the present invention. DETAILED DESCRIPTION

[0066] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. All other embodiments obtained by ordinary technicians in this field based on the embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0067] Reference Figure 1 As shown, the present invention discloses a method for evaluating the matching state of a turbine engine considering the swirl transfer of the air system, comprising the following steps:

[0068] S1. Obtain engine data and pre-process the data.

[0069] Furthermore, the engine data obtained include design point data, non-design point steady-state data and transition process transient data; from the data source point of view, it includes design data, simulation data and test measurement data.

[0070] Furthermore, the data should be preprocessed after collection.

[0071] Specifically, preprocessing should first eliminate the data of failed test points, that is, eliminate the data of test points that are obviously inconsistent with the magnitude or change trend of the measurement results of other test points.

[0072] Preprocessing should unify the data units. For example, in temperature data, the measurement results in degrees Celsius, commonly used in engineering, should be converted to data in the International System of Units (SI) of Kelvin. At the same time, the gauge pressure should be combined with the ambient pressure to convert it to absolute pressure.

[0073] Preprocessing involves filtering in the time dimension. Taking into account phenomena such as airflow pulsation within the disk cavity and the air ripples caused by the upstream blade wake sweeping across the downstream blade, the airflow pulsation data with a frequency above a given time threshold, which can be set at 5 Hz, is filtered.

[0074] Preprocessing also includes spatial averaging. For measurement parameters at the same lumped location in the flow direction, such as multiple measurement points arranged circumferentially on the same turbine disc, the measurement data should be averaged. For multiple measurement points arranged radially on the same turbine disc, the measurement data should be averaged, weighted by flow area or flow rate.

[0075] Furthermore, the node status parameters in the pre-processed turbine engine data include:

[0076] For main channel nodes, node state parameters include temperature, pressure, oil-gas ratio and water-gas ratio; for air system nodes, node state parameters include temperature, pressure, oil-gas ratio, water-gas ratio and swirl velocity; for thermal nodes, node state parameters include temperature; for mechanical nodes, node state parameters include speed; for components, they include flow rate, etc.

[0077] S2. Based on the engine data preprocessed in S1, considering the swirl transfer of the engine air system, a whole-machine transient model describing the aerodynamic thermodynamic matching conservation relationship in the engine is established.

[0078] S2.1. Model the turbine engine as a combination of nodes and elements. Figure 2 .

[0079] Furthermore, as a fluid-thermal machine, the turbine engine has node types including compressible fluid nodes, thermal nodes, and mechanical nodes. Among them, the compressible fluid nodes are divided into main channel fluid nodes and air system fluid nodes according to whether the swirl parameter transfer is considered.

[0080] Furthermore, the conservation equations for compressible fluid at the main channel nodes are as follows:

[0081]

[0082]

[0083] The above equations represent the conservation equations of mass, energy, water vapor component and oil and gas component of the compressible fluid. Among them, V represents the volume of the node, ρ represents the density of the compressible fluid, T represents the temperature of the compressible fluid, c v represents the specific heat ratio of compressible gas at constant volume, f fuel represents the oil and gas mass fraction of the compressible fluid, f water represents the water vapor mass fraction of the compressible fluid, τ represents time; n represents the number of elements connected to the node, Represents the mass flow rate into or out of a node; positive for flows into a node and negative for flows out of a node.

[0084] Furthermore, based on the conservation equations of the main channel nodes, the air system nodes also have the following angular momentum conservation relationship:

[0085]

[0086] Where r represents the average rotation radius at the node, V θ represents the angular velocity of the compressible fluid; it is defined as positive clockwise along the heading direction and negative counterclockwise. The above five conservation equations together constitute the aerodynamic thermodynamic conservation equations for compressible fluid nodes considering swirl transfer.

[0087] Furthermore, unlike the compressible fluid nodes or fluid domains mentioned above, the engine's structural components are solid domains. The heat transfer process within these solid domains interacts with the flow and heat exchange in the fluid domain, thus affecting the matching operating point of the entire engine. The transfer process in the solid domain of the engine can be modeled as a thermal node, at which the following conservation relations exist:

[0088]

[0089] The above formula describes the energy conservation relationship of the solid domain heat transfer process. p represents the constant-pressure specific heat of compressible gas, Q i is the node heat flow, which is positive when flowing into the node and negative when flowing out of the node.

[0090] Specifically, as a heat engine, an aircraft turbine engine has a power-generating end (e.g., turbine) and a power-consuming end (e.g., compressor and rotating disc cavity), which are connected by a shaft and transmit power. The power conservation relationship determines the rotor speed state, which in turn affects the working state of components such as the impeller machinery. Therefore, a mechanical node is introduced to implement the energy conservation relationship between the power-generating end and the power-consuming end at the mechanical node:

[0091]

[0092] The above equation describes the energy conservation relationship at the mechanical node. Where ω is the angular velocity of the rotor, J is the moment of inertia of the rotor, and W i Represents the transferred power. The power flowing into the node is positive, and the power flowing out of the node is negative.

[0093] The seven aerodynamic thermodynamic conservation equations above describe the conservation of flow, composition, angular momentum, and energy for the entire engine under transient conditions. They serve as the governing equations for the nodes in the transient model of an aircraft turbine engine. Solving this system of equations enables the conservation and transfer of swirl in the air system and the determination of the matching operating point for the entire engine. By setting the transient terms on the left side of each of the above equations to zero, the governing equations for the nodes in the steady-state engine are obtained.

[0094] Furthermore, corresponding to the nodes, the component module is mainly used to describe the response of engine components under the excitation of the node state. The component module mainly includes the main channel component, the inlet / exhaust component and the air system component.

[0095] Specifically, the main flow channel components include the inlet, compressor, combustion chamber, turbine, and tailpipe. Main flow channel components, disregarding swirl effects, are connected to the main flow channel nodes. Their inputs are the state parameters (i.e., pressure, temperature, oil-to-air ratio, and water-to-air ratio) at the upstream and downstream main flow channel nodes and the state parameters (i.e., rotational speed) at the mechanical nodes. Their outputs are the flow rate, enthalpy flow, oil-to-air flow rate, and water-to-air flow rate at the element's inlet and outlet. This response depends on the main flow channel component's characteristic diagrams, such as the turbomechanical pressure ratio and efficiency characteristics.

[0096] Specifically, the air system should include flow paths that perform the main functions and have large flow rates, such as turbine disk cooling and sealing flow paths, turbine outer ring cooling flow paths, turbine blade air supply flow paths, etc. The same flow path should include components that have an important influence on the pressure and temperature distribution along the process, such as pre-swirl nozzles, disc cavities, grate teeth, and vortex reducers. Since swirl transfer is considered in the air system, the product of the circumferential velocity and the radial position, that is, the circulation, is actually transmitted using mass flow as the carrier. This physical quantity is correlated with the radial position. Therefore, the outlet of the upstream element and the inlet of the downstream element should have the same radial position, that is, share the same node. The node should cover the key sections or cavity positions of concern, and be appropriately divided considering the actual measurement point positions. The volume inertia and radius parameters should be set at the nodes.

[0097] Furthermore, the air system components need to return mass flow, enthalpy flow, oil and gas mass flow, water vapor mass flow and angular momentum flow to the upstream and downstream nodes under the boundary conditions set by the upstream and downstream nodes (such as pressure, temperature, oil-gas ratio, water vapor ratio and swirl velocity). Among them, the mass flow and enthalpy flow can be obtained by characteristic diagram interpolation method or theoretical methods such as isentropic expansion relationship and radial equilibrium relationship. On this basis, according to the boundary integral method, the physical quantity returned to the upstream node can be simply obtained. When solving the physical quantity returned by the component to the downstream node, in order to consider the transmission of the swirl coefficient, it is necessary to consider the influence of windage torque and temperature rise. The control equation of the air system component is:

[0098]

[0099] The above formula represents the flow rate, windage temperature rise, composition, and swirl relationship of the inlet and outlet of the component. Among them, the subscripts in and out represent the inlet and outlet of the component respectively; represents the mass flow rate of the fluid flowing through the element, T represents the temperature of the compressible fluid, c p represents the constant-pressure specific heat of compressible gas, f fuel represents the oil and gas mass fraction of the compressible fluid, fwater represents the water vapor mass fraction of the compressible fluid, r represents the average rotation radius of the element, V θ represents the angular velocity of the compressible fluid; M rotor ω represents the windage torque, Q c Represents the heat exchange between the component and the thermal node; M rotor ,M stator Representing the torques exerted by the rotor and stator on the airflow, respectively, they can generally be derived from empirical relationships, with circumferential acceleration of the fluid considered positive and deceleration negative. It is important to note that when calculating the change in angular momentum, the torques exerted by both the rotor and stator on the fluid must be considered. When calculating temperature rise, since the stator's displacement is zero in the stationary coordinate system, only the windage temperature rise effect of the rotor needs to be considered.

[0100] Specifically, the bleed / exhaust element connects the main flow path node and the air system node, automatically calculates the air system bleed / exhaust parameters, and provides a boundary for air system component calculations. The bleed / exhaust element primarily includes turbomachinery components such as compressor bleed air and turbine confluence. The total temperature, total pressure, static pressure, and swirl velocity at the bleed / exhaust location are automatically calculated using the following method.

[0101] Furthermore, we first define the enthalpy rise ratio coefficient of the turbomachinery, which is used to describe the relationship between the total enthalpy at the inlet / exhaust position and the total enthalpy at the compressor inlet:

[0102]

[0103] Among them, H in ,H out ,H bleed The separate tables represent the total enthalpy of the fluid at the inlet, outlet, and inlet / exhaust positions of the impeller machinery; α represents the enthalpy rise ratio coefficient, which can generally be regarded as a constant under normal operating conditions. Therefore, based on the inlet temperature, inlet pressure, and outlet pressure of the impeller machinery, the total enthalpy at the outlet of the impeller machinery can be obtained by interpolation of the characteristic diagram, and then the total enthalpy at the inlet / exhaust position of the impeller machinery can be obtained based on the enthalpy rise ratio coefficient:

[0104] H bleed =α·(H out -H in )+H in

[0105] On this basis, the total temperature at the inlet / exhaust position can be calculated. The total pressure at the inlet / exhaust position of the impeller machine can be calculated through the adiabatic process:

[0106]

[0107] Among them, P t,in ,T t,in ,Pt,bleed ,T t,bleed are the total pressure and total temperature at the inlet and exhaust / intake positions of the turbomachinery, respectively. γ represents the polynomial index of this thermodynamic process, which can be inversely solved through the total pressure ratio and efficiency of the turbomachinery.

[0108] Furthermore, since the air system's inlet / exhaust parameters are usually total temperature and static pressure, it is also necessary to calculate the static pressure at the inlet / exhaust position. Based on the aerodynamic thermodynamic function, the axial velocity coefficient λ at the inlet / exhaust position can be calculated: axial :

[0109]

[0110] Among them, q(λ axial ) is the dimensionless density flow of the main channel gas, is the mass flow rate of the gas in the main channel, A is the effective flow area of ​​the gas channel, K is a constant related to the specific heat ratio k of the gas; P t,bleed ,T t,bleed are the total pressure and total temperature at the turbine machine inlet / exhaust positions respectively;

[0111] Therefore, the axial velocity coefficient λ of the gas at the inlet / exhaust can be obtained by inversely solving the aerodynamic thermodynamic function. axial .

[0112] Since the direction of the airflow inside the turbomachinery is not always consistent with the engine axis, especially at the inlet of the compressor blades and the outlet of the turbine blades, the circumferential velocity of the airflow is relatively high, which may significantly affect the total flow velocity of the airflow and must be taken into account. Assuming that under normal operating conditions, the airflow does not significantly separate when it flows out of the blades, the airflow angle relative to the blade structure can be ignored. The equivalent airflow angle can be approximated by the angle between the outlet position of the blade structure and the engine axis. Therefore, the circumferential velocity coefficient at this position relative to the blade can be calculated using trigonometric functions:

[0113] λ θ,vane =λ axial tan(θ)

[0114] Among them, λ θ,vane ,λ axial are the relative circumferential velocity coefficient and axial velocity coefficient respectively, and θ is the angle between the outlet position of the blade structure and the engine axis.

[0115] For the airflow angle behind the guide vane, the swirl coefficient in the stationary coordinate system is λ θ，vane , and for the rotating rotor blades, it is also necessary to transform the coordinate system to obtain the swirl velocity in the stationary coordinate system:

[0116]

[0117] Among them, λ θ,blade ,λ θ,vane represent the absolute circumferential velocity coefficient and the relative circumferential velocity coefficient respectively; is the rotor rotation speed coefficient, ω is the rotor speed, r is the radius at the inlet / exhaust position, R is the gas constant, k is the specific heat ratio of the gas, T t,bleed are the total temperatures at the turbine inlet and outlet locations, respectively.

[0118] The swirl velocity at the inlet / exhaust position can be calculated:

[0119]

[0120] Among them, V θ ,λ θ are the circumferential velocity and circumferential velocity coefficient of the gas, R is the gas constant, k is the specific heat ratio of the gas, T t,bleed are the total temperatures at the turbine inlet and outlet locations, respectively.

[0121] The total velocity coefficient at the inlet / exhaust position is the composite velocity coefficient of the axial and circumferential directions:

[0122]

[0123] Among them, λ bleed ,λ θ ,λ axial They are the total velocity coefficient, circumferential velocity coefficient and axial velocity coefficient at the inlet / exhaust position respectively.

[0124] The static pressure at the inlet / exhaust position can be obtained by the aerodynamic thermodynamic relationship:

[0125] P s,bleed =P t,bleed ·π(λ bleed )

[0126] Among them, P s,bleed ,P t,bleed are the static pressure and total pressure at the inlet / exhaust position, respectively, bleed is the total velocity coefficient at the inlet / exhaust position, π(λ bleed ) is the pressure ratio function.

[0127] Furthermore, for the confluence position, the parameters of the mixed main channel airflow and the air system airflow at the inlet / exhaust position can be obtained by the flow weighted average method, which is used to solve the main channel components.

[0128] Therefore, the total temperature, static pressure and swirl velocity at the inlet / exhaust position can be automatically calculated by the pressure ratio, efficiency, enthalpy rise ratio coefficient, gas channel flow area and blade structure angle or airflow angle of the impeller machinery for the calculation of air system components.

[0129] In summary, the entire aircraft turbine engine can be modeled into node modules (including main channel compressible fluid nodes, air system compressible fluid nodes, thermal nodes, and mechanical nodes) and component modules (main channel components, inlet / exhaust components), to achieve global conservation of the main channel and the air system, automatically match the inlet / exhaust parameters of the main channel and the air system at the interface with the main channel, and realize the transfer of swirl parameters between the various components within the air system, see Figure 3 .

[0130] S2.2. Integrate the main flow channel components, air system components considering swirl, and inlet / exhaust components, adjust the component parameters based on the preprocessed data, and obtain the transient model of the entire machine.

[0131] Specifically, the main channel elements are connected through the main channel fluid nodes, and the air system elements are connected through the air system fluid nodes; the main channel elements and the air system are connected through the main channel fluid nodes, and the intake / exhaust elements and the air system fluid nodes; the elements with power generation or consumption are connected through mechanical nodes, and the elements with heat conduction and transfer are connected through thermal nodes.

[0132] The model solution is controlled by parameter handles such as the pressure ratio scaling factor, flow scaling factor, and efficiency scaling factor of the main channel element, the enthalpy rise ratio of the exhaust / inlet element, and the flow scaling factor and wind resistance scaling factor of the air system element. The relative error between the partial solution results and the relevant values ​​in the preprocessed data is within an acceptable range, and the transient model of the entire machine can be obtained.

[0133] The calculation of the relative error between the model and the data should take into account multiple engine states or the entire transition process, and the model parameters should be adjusted based on criteria such as least squares.

[0134] S3. Solve the whole machine transient model established in S2, obtain the engine operating state parameters, and realize the engine matching state evaluation.

[0135] Specifically, the boundary conditions under the working conditions to be evaluated are input into the engine model, including the engine flight altitude, Mach number, and fuel flow rate; by solving the engine model under the working conditions to be evaluated, the matching state parameters of the engine under the working conditions can be obtained, which can be used to evaluate the performance and safety level of the engine design.

[0136] The model solution method can be a gradient-driven algorithm such as the secant method, or a target-driven algorithm such as the control algorithm; the matching state parameters obtained include engine thrust, engine rotor speed, engine rotor axial force, turbine inlet temperature, engine exhaust temperature, turbine disc cavity temperature and compressor exhaust pressure.

[0137] Based on the method proposed in this paper, a transient model of a turboshaft engine was established, including the main flow channel and air system components. Figure 4 The inlet / exhaust parameters at the compressor, gas turbine, and power turbine were automatically calculated using the proposed method. The inlet / exhaust parameters were compared with the turbomachinery three-dimensional calculation method, and the results are shown in Table 1.

[0138] Table 1

[0139]

[0140] It can be seen that the proposed method has good accuracy under multiple stable engine states. Furthermore, during the engine transient process, the relative errors of key cross-sectional parameters of the main flow channel and key cavity temperature and cavity pressure parameters of the air system compared with experimental values ​​are all less than 5%. This demonstrates the high computational accuracy of the proposed method.

[0141] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. Reference can be made to the common and similar parts between the various embodiments. For the devices disclosed in the embodiments, since they correspond to the methods disclosed in the embodiments, the description is relatively simple, and the relevant parts can be referred to the method description.

[0142] The above description of the disclosed embodiments is intended to enable one skilled in the art to implement or use the present invention. Various modifications to these embodiments will be readily apparent to one skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the present invention. Therefore, the present invention is not limited to the embodiments shown herein but is intended to conform to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A method for evaluating the matching status of a turbine engine considering the swirl transfer of the air system, characterized in that: The following steps are involved: S1. Obtain engine data and pre-process the engine data; S2. Based on the engine data preprocessed in S1, considering the swirl transfer of the engine air system, a whole-machine transient model describing the aerodynamic thermodynamic matching conservation relationship in the engine is established; S3, solving the whole machine transient model established in S2, obtaining the engine operating state parameters, and realizing the engine matching state evaluation; The turbine engine model is established in S2 as follows: The turbine engine is modeled as a combination of nodes and elements. As a fluid-thermal machine, the turbine engine has node types including compressible fluid nodes, thermal nodes, and mechanical nodes. Compressible fluid nodes are divided into two types: main channel nodes and air system nodes, depending on whether swirl parameter transfer is considered. Elements include main channel elements, inlet / exhaust elements, and air system elements. The conservation equations for the mass, energy, water vapor component, and oil and gas component of the compressible fluid at the main channel node are as follows: Where V represents the volume of the node, ρ represents the density of the compressible fluid, T represents the temperature of the compressible fluid, c v represents the specific heat ratio of compressible gas at constant volume, f fuel represents the oil and gas mass fraction of the compressible fluid, f water represents the water vapor mass fraction of the compressible fluid, τ represents time; n represents the number of elements connected to the node, Represents the mass flow into or out of a node.

2. A turbine engine matching status assessment method considering air system swirl transfer according to claim 1, characterized in that: Based on the conservation equations of the main channel nodes, the air system nodes also have the following angular momentum conservation relationship: Where r represents the average rotation radius at the node, V θ represents the angular velocity of the compressible fluid.

3. The method for evaluating the matching status of a turbine engine considering swirl transfer in an air system according to claim 1, characterized in that: The transfer process in the solid domain of the engine is modeled as a thermal node, where the following energy conservation relationship exists: Among them, c p represents the constant-pressure specific heat of compressible gas, Q i is the node heat flow, which is positive when flowing into the node and negative when flowing out of the node.

4. The method for evaluating the matching status of a turbine engine considering swirl transfer in an air system according to claim 1, characterized in that: The energy conservation relationship between the power generation end and the power consumption end at the mechanical node is as follows: Where ω is the angular velocity of the rotor, J is the moment of inertia of the rotor, and W i Represents the transferred power.

5. The method for evaluating the matching status of a turbine engine considering swirl transfer in an air system according to claim 1, characterized in that: The governing equations for air system components represent the flow rate, windage temperature rise, composition, and swirl relationships at the inlet and outlet of the components, respectively: Among them, the subscripts in and out represent the inlet and outlet of the component respectively; represents the mass flow rate of the fluid flowing through the element, T represents the temperature of the compressible fluid, c p represents the constant-pressure specific heat of compressible gas, f fuel represents the oil and gas mass fraction of the compressible fluid, f water represents the water vapor mass fraction of the compressible fluid, r represents the average rotation radius of the element, V θ represents the angular velocity of the compressible fluid; M rotor ω represents the windage torque, Q c Represents the heat exchange between the component and the thermal node; M rotor ,M stator represent the torques exerted by the rotor and stator on the airflow respectively.

6. The method for evaluating the matching status of a turbine engine considering swirl transfer in an air system according to claim 1, characterized in that: First, define the enthalpy rise ratio coefficient of the impeller machine: Among them, H in ,H out ,H bleed The separate tables represent the total enthalpy of the fluid at the inlet, outlet, and exhaust positions of the impeller machinery. Based on the inlet temperature, inlet pressure, and outlet pressure of the impeller machinery, the total enthalpy at the outlet of the impeller machinery is obtained by interpolation of the characteristic diagram, and then the total enthalpy at the exhaust position of the impeller machinery is obtained based on the enthalpy rise ratio coefficient: H bleed =α·(H out -H in )+H in The total pressure at the impeller machine inlet / exhaust position is calculated through the adiabatic process: Among them, P t,in ,T t,in ,P t,bleed ,T t,bleed are the total pressure and total temperature at the turbomachinery inlet and introduction / exhaust positions, respectively, and γ represents the polynomial index of this thermodynamic process.

7. The method for evaluating the matching status of a turbine engine considering swirl transfer in an air system according to claim 6, characterized in that: Based on the aerodynamic thermodynamic function, the axial velocity coefficient λ at the inlet / exhaust position is calculated axial : Among them, q(λ axial ) is the dimensionless density flow of the main channel gas, is the mass flow rate of the gas in the main channel, A is the effective flow area of ​​the gas channel, K is a constant related to the specific heat ratio k of the gas; P t,bleed ,T t,bleed are the total pressure and total temperature at the turbine machine inlet / exhaust positions respectively; The circumferential velocity coefficient is obtained by trigonometric function: l θ,vane =λ axial ·tan(θ) Among them, λ θ,vane ,λ axial are the relative circumferential velocity coefficient and axial velocity coefficient respectively, and θ is the angle between the outlet position of the blade structure and the engine axis; For the airflow angle behind the guide vane, the swirl coefficient in the stationary coordinate system is λ θ,vane , and for the rotating rotor blades, the coordinate system is transformed to obtain the swirl velocity in the stationary coordinate system: Among them, λ θ,blade ,λ θ,vane represent the absolute circumferential velocity coefficient and the relative circumferential velocity coefficient respectively; is the rotor rotation speed coefficient, ω is the rotor speed, r is the radius at the inlet / exhaust position, R is the gas constant, k is the specific heat ratio of the gas, T t,bleed are the total temperatures at the turbine inlet / exhaust locations respectively; Calculate the swirl velocity at the inlet / exhaust position: Among them, V θ ,λ θ are the circumferential velocity and circumferential velocity coefficient of the gas, R is the gas constant, k is the specific heat ratio of the gas, T t,bleed are the total temperatures at the turbine inlet / exhaust locations respectively; The total velocity coefficient at the inlet / exhaust position is the composite velocity coefficient of the axial and circumferential directions: Among them, λ bleed ,λ θ ,λ axial are the total velocity coefficient, circumferential velocity coefficient and axial velocity coefficient at the inlet / exhaust position respectively; The static pressure at the inlet / exhaust position is obtained by the aerodynamic thermodynamic relationship: P s,bleed =P t,bleed ·p(l bleed ) Among them, P s,bleed ,P t,bleed are the static pressure and total pressure at the inlet / exhaust position, respectively, bleed is the total velocity coefficient at the inlet / exhaust position, π(λ bleed ) is the pressure ratio function.

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