A method and system for optimizing cycle parameters of an aviation gas turbine shaft engine

Abstract

The present invention discloses a cycle parameter optimization method and system for an aviation gas turbine shaft engine, the method comprising: based on helicopter engine system basic data, selecting engine cycle parameters to be optimized and their corresponding ranges; determining the power unit mass of the helicopter according to the helicopter engine system basic data and the engine cycle parameters to be optimized; simulating the helicopter flight trajectory based on the helicopter engine system basic data to obtain the helicopter flight fuel mass; using the helicopter power unit mass and the flight fuel mass to obtain the helicopter take-off mass and perform a test, and calculating the helicopter efficiency criterion after the test passes; determining the optimization optional domain of the engine cycle parameters to be optimized according to the calculation result of the helicopter efficiency criterion. Considering the helicopter efficiency index in the preliminary design stage of the engine is conducive to improving the adaptability of the engine and the helicopter.

Description

Technical Field

[0001] The present invention relates to the technical field of aviation gas turbine shaft engines, and in particular to a cycle parameter optimization method and system for aviation gas turbine shaft engines. Background Art

[0002] The power unit of modern helicopters is generally provided by aviation gas turbine shaft engines to provide power for helicopters to meet the power requirements of various flight states of helicopters. For aviation gas turbine shaft engines, the flight conditions of helicopters have a great influence on the performance of the engine; when the air temperature rises or the altitude increases, the performance of the engine will drop rapidly, resulting in a sharp deterioration in the helicopter's carrying capacity and maneuverability, and an increase in fuel consumption. Therefore, under this condition, the matching of the engine and the helicopter rotor system requires the engine to have a higher power, and at the same time, the engine power must be limited when flying at low temperatures close to the ground.

[0003] It can be seen that helicopters have higher requirements for engine selection, not only requiring the engine to meet the basic conditions of the helicopter's power requirements, but also requiring the engine's power state to match the type and purpose of the helicopter, and improve the helicopter's efficiency index. However, the use of an oversized engine with power restrictions will weaken the helicopter's characteristics and rapidly increase the engine's mass.

[0004] In order to solve this problem, it is necessary to consider the normal and extreme conditions of use according to the helicopter type and purpose in the process of determining the cycle parameters at the beginning of engine design. Even if only a part of it is considered, it will help solve the problem to a certain extent. However, the starting point of the existing work on determining the cycle parameters of aviation gas turbine shaft engines only considers the performance indicators of the engine, such as unit power, fuel consumption rate, etc. Summary of the invention

[0005] The present invention aims to provide a method and system for optimizing the cycle parameters of an aviation gas turbine shaft engine, so as to take into account the helicopter efficiency index in the preliminary design stage of the engine, which is beneficial to improving the compatibility of the engine and the helicopter.

[0006] To achieve the above object, the present invention provides a method for optimizing cycle parameters of an aviation gas turbine shaft engine, comprising:

[0007] Based on the basic data of the helicopter-engine system, select the engine cycle parameters to be optimized and their corresponding ranges;

[0008] Determining the power unit mass of the helicopter based on the helicopter-engine system basic data and the engine cycle parameters to be optimized;

[0009] Based on the basic data of the helicopter-engine system, simulate the flight trajectory of the helicopter to obtain the flight fuel quality of the helicopter;

[0010] Using the helicopter's power unit mass and flight fuel mass, obtain the helicopter's takeoff mass and conduct a test. After the test is passed, calculate the helicopter's efficiency criterion;

[0011] According to the calculation results of the helicopter efficiency criterion, the optimization optional domain of the engine cycle parameters to be optimized is determined.

[0012] Furthermore, the selecting of engine cycle parameters to be optimized based on the helicopter-engine system basic data includes:

[0013] The helicopter-engine system basic data includes: engine basic data, helicopter basic data, use condition data and engine cycle parameters to be optimized;

[0014] The engine cycle parameters to be optimized include compressor pressure ratio and combustion chamber outlet temperature.

[0015] Further, the determining the power unit mass of the helicopter according to the helicopter-engine system basic data and the engine cycle parameters to be optimized includes:

[0016] Obtaining flow channel dimensions of engine components according to the helicopter-engine system basic data and the engine cycle parameters to be optimized;

[0017] Determine the engine mass based on the relationship between the flow channel size of the engine components and the mass of the corresponding components;

[0018] The power unit mass is obtained according to the engine mass.

[0019] Furthermore, the obtaining of the flow channel dimensions of the engine components according to the helicopter-engine system basic data and the engine cycle parameters to be optimized includes:

[0020] According to the basic data of the helicopter-engine system and the engine cycle parameters to be optimized, the inlet and outlet cross-sectional parameters and performance parameters of each engine component are calculated using the engine aerodynamic thermodynamic calculation mathematical model;

[0021] Determine the flow channel dimensions of engine components based on the inlet and outlet cross-sectional parameters and performance parameters of each engine component;

[0022] Among them, the engine components include: air intake, compressor, gas turbine, combustion chamber, power turbine, casing and transition section.

[0023] Further, obtaining the power unit mass according to the engine mass includes:

[0024] Determining the mass of the power unit based on the mass of the engine, the mass of the components of the conventional installation system of the gas turbine engine on the helicopter, the mass of the components of the cooling and fire protection system, and the mass of the components of the dust removal system;

[0025] If the engine does not have a reducer, the mass of the power unit is:

[0026] M cy =n eng K cy M eng

[0027] Where M cy Indicates the mass of the helicopter's power unit when the engine is not equipped with a reducer; n eng Indicates the number of engines installed on the helicopter; M eng Indicates the mass of a single engine; K cy It represents the engine mass increase factor caused by considering the conventional installation system components, cooling and fire protection system components and dust removal system components of the gas turbine engine on the helicopter;

[0028] If the engine is equipped with a reducer, the mass of the power unit is:

[0029] M′ cy =n eng (K cy M' eng +M red +M sys )

[0030]

[0031] In the formula, M′ cy Indicates the weight of the helicopter's power plant when the engine is equipped with a reducer; M red Indicates the mass of the engine reducer; PW e represents the design input power of the reducer; n2 represents the design input speed of the reducer; M sys Indicates the mass of the power unit accessories and system (this mass remains unchanged when the engine mass changes); M' eng Indicates the engine mass without reduction gear and accessory system.

[0032] Furthermore, simulating the helicopter flight trajectory based on the helicopter-engine system basic data to obtain the helicopter flight fuel quality includes:

[0033] Based on the helicopter-engine system basic data, dividing the helicopter flight trajectory into a plurality of flight trajectory segments according to the helicopter flight mission profile;

[0034] The helicopter aerodynamic performance calculation mathematical model is used to calculate the required power for each section of the helicopter flight trajectory;

[0035] Calculate the engine fuel consumption rate corresponding to the required power by using the engine aerodynamic thermodynamic calculation mathematical model;

[0036] The helicopter flight fuel mass is obtained according to the required power and its corresponding engine fuel consumption rate.

[0037] Furthermore, the helicopter flight fuel mass is:

[0038]

[0039] Among them, M mL represents the helicopter flight fuel mass applicable to all helicopters, t n Indicates the total flight time of the helicopter to complete the flight; M m represents the mass of fuel consumed; r represents the total number of flight trajectory segments; i represents the number of flight trajectory segments; Δt i PW represents the flight time on the i-th flight track segment; i With sfc i They represent the required power and fuel consumption rate for the i-th flight segment respectively;

[0040] If the helicopter flight trajectory is a simple take-off-cruise at a certain altitude-landing type continuous cruise flight segment, the helicopter flight fuel mass is:

[0041] M' mL =n eng k L sfc P PW P L L / V L

[0042] Among them, M′ mL Indicates the helicopter flight fuel mass when the helicopter flight trajectory is the continuous cruising flight segment; n eng Indicates the number of engines installed on the helicopter; PW p With sfc p They represent the average power and fuel consumption rate of the helicopter during the cruising flight phase; k L It represents the coefficient that considers the influence of fuel flow rate in other flight stages other than the cruise stage; L L Indicates the helicopter range; V L Indicates the helicopter flight technical speed, which is the average speed considering all flight phases, based on the cruising speed V p Value determination: V L =k V Vp , k V Indicates the correction factor.

[0043] Furthermore, before obtaining the helicopter take-off mass by using the power plant mass and the helicopter flight fuel mass, a first tolerance coefficient ε is used. m Determine the helicopter flight fuel mass M mL or M′ mL Whether the initial given flight fuel mass M is met mLs :

[0044] When the calculated flight fuel mass of the helicopter meets the initially given flight fuel mass, the helicopter take-off mass is calculated and obtained by using the helicopter power unit mass and the helicopter flight fuel mass;

[0045] When the calculated flight fuel mass of the helicopter does not meet the initially given flight fuel mass, reselecting the value of the engine cycle parameter to be optimized within the initially given range of the optimized engine cycle parameter, and recalculating the helicopter power unit mass and the helicopter flight fuel mass until the helicopter flight fuel mass meets the initially given flight fuel mass;

[0046] Among them, if or , the calculated flight fuel mass of the helicopter meets the initially given flight fuel mass.

[0047] Furthermore, the helicopter take-off mass is obtained and tested by using the helicopter power unit mass and the helicopter flight fuel mass. After the test is passed, the helicopter efficiency criterion is calculated, including:

[0048] The helicopter take-off mass is:

[0049] M oc =M load +M ce +M enp +M me

[0050] Among them, M oc represents the take-off mass of the helicopter; M load Represents the effective load of the helicopter, which is given when the helicopter is designed; M enp Indicates the mass of the helicopter excluding the power unit and fuel system, including the mass of the helicopter body, equipment and crew; M ce The mass M of the helicopter power unit when the engine is not equipped with a reducer cy Or the helicopter power unit mass M' when the engine is equipped with a reducer cy ;M meDenotes the helicopter flight fuel mass M applicable to all helicopters mL Or the helicopter flight fuel mass M′ when the helicopter flight trajectory is the continuous cruise flight segment mL ;

[0051] Using the second tolerance factor ε m0 , determine the helicopter take-off mass M oc Whether the initial take-off mass M0 is met:

[0052] If the helicopter take-off mass M oc If the initial given take-off mass M0 is satisfied, the helicopter efficiency criterion is calculated;

[0053] If the helicopter take-off mass M oc If the initial given takeoff mass M0 is not satisfied, the value of the engine cycle parameter to be optimized is reselected within the initial given range of the optimized engine cycle parameter, and the helicopter power unit mass and the helicopter flight fuel mass are calculated again until the helicopter takeoff mass M0 is satisfied. oc Meet the initial given take-off mass M0;

[0054] Among them, when When the helicopter take-off mass M oc Meet the initial given take-off mass M0.

[0055] Furthermore, according to the calculation results of the helicopter efficiency criterion, the optimization optional domain of the engine cycle parameters to be optimized is determined, including:

[0056] According to the purpose and flight status characteristics of the helicopter, one or more criteria are selected for calculation to obtain corresponding calculation values;

[0057] According to the calculated values ​​of one or more criteria in the selected helicopter efficiency criteria, the extreme value and compatible range of each criterion are correspondingly determined;

[0058] The calculated values ​​based on one or more criteria, as well as the extreme values ​​and compatible ranges corresponding to each criterion, will satisfy The value range of the engine cycle parameter to be optimized is used as the feasible domain of parameter optimization of the helicopter efficiency criterion; wherein, Y i , Y opt,i , ΔY i They represent the calculated value, extreme value and compatible range of the i-th criterion respectively;

[0059] Determining an optimization optional domain of the engine cycle parameters to be optimized according to the parameter optimization feasible domain of the helicopter efficiency criterion;

[0060] The helicopter efficiency criteria include a power unit efficiency criterion, a transport capacity criterion, a transport efficiency criterion and a fuel consumption rate criterion.

[0061] Further, according to the mass of the power plant and the mass of the helicopter flight fuel, the efficiency criterion of the power plant is obtained;

[0062] The power plant efficiency criterion is:

[0063]

[0064] Among them, γ e Indicates the power plant efficiency criterion; M ce The mass M of the helicopter power unit when the engine is not equipped with a reducer cy Or the helicopter power unit mass M' when the engine is equipped with a reducer cy ;M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or when the helicopter flight trajectory is the continuous cruise flight segment, the helicopter flight fuel mass M′ mL ;n eng Indicates the number of engines installed on the helicopter; PW indicates the power of a single engine; K cy Represents the term related to the engine power-to-weight ratio; K sfc Indicates items related to engine fuel consumption rate;

[0065] The transport capacity criteria are:

[0066]

[0067] Among them, γ load Indicates the transport capacity criterion; M load Indicates the helicopter's payload; M emp Indicates the unladen mass of the helicopter; V L It represents the flight technical speed, which is the average speed considering all flight phases, according to the cruising speed V p Value determination: V L =k V V p , k V represents the correction factor;

[0068] The transport efficiency criterion is:

[0069]

[0070] Among them, γ leff represents the transport efficiency criterion; k L V represents the coefficient for considering the influence of fuel flow rate in other flight stages other than the cruise stage; PIndicates cruising speed; M cn Indicates the helicopter crew mass; L L Indicates helicopter range; M emp Indicates the unladen mass of the helicopter;

[0071] The fuel consumption rate criteria are:

[0072]

[0073] Among them, γml represents the fuel consumption rate criterion; ε represents the effective load utilization rate.

[0074] Based on the same inventive concept, the present invention also provides a cycle parameter optimization system for an aviation gas turbine shaft engine, comprising a selection unit, an acquisition unit, a calculation unit and a determination unit.

[0075] The selection unit is used to select the engine cycle parameters to be optimized and their corresponding ranges based on the helicopter-engine system basic data;

[0076] The acquisition unit is used to calculate and acquire the mass of the helicopter's power plant according to the helicopter-engine system basic data and the engine cycle parameters to be optimized; and is also used to simulate the helicopter's flight trajectory based on the helicopter-engine system basic data to acquire the helicopter's flight fuel mass;

[0077] The calculation unit is used to obtain the helicopter take-off mass by using the helicopter's power plant mass and flight fuel mass, and to check the helicopter's take-off mass, and to calculate the helicopter efficiency criterion after the check is passed;

[0078] The determination unit is used to determine the optimization optional domain of the engine cycle parameter to be optimized according to the calculation result of the helicopter efficiency criterion.

[0079] Based on the same inventive concept, an embodiment of the present invention also provides an electronic device, comprising: a memory, a processor, and a computer program stored in the memory and running on the processor, wherein when the processor executes the computer program, it implements the aforementioned method for optimizing the cycle parameters of an aviation gas turbine shaft engine.

[0080] Based on the same inventive concept, an embodiment of the present invention further provides a computer storage medium, wherein the computer storage medium stores computer executable instructions, and when the computer executable instructions are executed, the aforementioned method for optimizing cycle parameters of an aviation gas turbine shaft engine is implemented.

[0081] The technical effects and advantages of the present invention are as follows: in the process of optimizing the engine cycle parameters by the method of the present invention, the purpose of the helicopter, the multi-state nature of the engine when used in the helicopter, and the designed range of the helicopter can be fully considered; the helicopter efficiency index is also taken as the goal of the cycle parameter optimization. Through the method of the present invention, the helicopter efficiency index can be considered in the preliminary design stage of the engine, which is beneficial to improving the compatibility of the engine and the helicopter.

[0082] Other features and advantages of the present invention will be described in the following description, and partly become apparent from the description, or understood by practicing the present invention. The purpose and other advantages of the present invention can be realized and obtained by the structures pointed out in the description, claims and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0083] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for use in the description of the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0084] Figure 1 This is a flow chart of a method for optimizing cycle parameters of an aviation gas turbine shaft engine according to an embodiment of the present invention;

[0085] Figure 2 A detailed flow chart of the steps of a method for optimizing cycle parameters of an aviation gas turbine shaft engine according to an embodiment of the present invention;

[0086] Figure 3 A schematic diagram of a feasible region for optimizing engine cycle parameters in an embodiment of the present invention;

[0087] Figure 4 This is a schematic structural diagram of a cycle parameter optimization system for an aviation gas turbine shaft engine according to an embodiment of the present invention;

[0088] Figure 5 The figure is a schematic diagram of the structure of an electronic device in an embodiment of the present invention. DETAILED DESCRIPTION

[0089] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0090] The aviation gas turbine shaft engine proposed by the present invention includes a single-rotor gas generator turboshaft engine, a single-rotor gas generator turboshaft engine with a free turbine, and a double-rotor gas generator turboshaft engine with a free turbine.

[0091] To address the deficiencies of the prior art, the present invention discloses a method for optimizing cycle parameters of an aviation gas turbine shaft engine. Figure 1 and Figure 2 As shown, the following steps are included:

[0092] Step 1: Based on the basic data of the helicopter-engine system, select the engine cycle parameters to be optimized and their corresponding ranges; specifically including:

[0093] The helicopter-engine system basic data includes: engine basic data, helicopter basic data, use condition data and engine cycle parameters;

[0094] The basic engine data include: efficiency indicating the technical level of compressor, gas turbine and power turbine, standard characteristic curves of compressor, gas turbine and power turbine, combustion efficiency and pressure loss coefficient of combustion chamber, relationship curve between cooling air volume and combustion chamber outlet temperature, pressure loss curves of inlet duct, tail nozzle and transition section, engine life index, material strength and cycle performance, and other parameters for engine cycle analysis.

[0095] The basic data of the helicopter include: rotor aerodynamic performance parameters, tail rotor aerodynamic performance parameters, fuselage drag coefficient and lift coefficient, mass excluding power unit and fuel, and helicopter take-off mass.

[0096] The operating conditions data include: flight profile, design mission (payload, range, etc.), technical requirements (helicopter dynamic ceiling, static ceiling, maximum cruising speed, etc.), engine usage restrictions, and helicopter usage restrictions.

[0097] Among the engine cycle parameters, the engine cycle parameters to be optimized and their corresponding ranges are selected; the ranges of the engine cycle parameters are determined according to the power required by the helicopter and the selected thermal parameter route; wherein the engine cycle parameters to be optimized include the compressor pressure ratio π k and the combustion chamber outlet temperature T t4 ;

[0098] The engine cycle parameters to be optimized also include the low compressor pressure ratio and the high compressor pressure ratio of the dual-rotor gas generator.

[0099] Step 2, determining the power unit mass of the helicopter according to the basic data of the helicopter-engine system and the engine cycle parameters to be optimized; comprising:

[0100] Based on the basic engine data in the helicopter-engine system basic data and the engine cycle parameters to be optimized, the engine aerodynamic and thermal calculation mathematical model is obtained; then, the engine aerodynamic and thermal calculation mathematical model based on the component method is used to calculate the engine design point and non-design point performance to obtain the inlet and outlet cross-sectional parameters and performance parameters of each engine component;

[0101] In this stage, the characteristics of components such as compressors, gas turbines, and power turbines are scaled using standard characteristics or characteristics of reference objects.

[0102] Based on the inlet and outlet cross-sectional parameters and performance parameters of each engine component, and taking into account the engine design technical requirements and constraints, the flow channel dimensions of the engine components are calculated and obtained respectively, and finally the flow channel dimensions of the entire engine are obtained.

[0103] The engine components include: air intake device or particle separator, compressor, gas turbine, combustion chamber, power turbine, casing, transition section and tail nozzle.

[0104] The engine mass is calculated and obtained based on the relationship between the flow path size of the engine components and the mass of the corresponding components, as well as the flow path size of the entire engine; or based on the existing engine mass statistical data, the correlation between the engine mass and its operating parameters is established to calculate the engine mass.

[0105] The mass of the power unit is obtained by calculation based on the mass of the engine, the mass of the components of the conventional installation system of the gas turbine engine on the helicopter, the mass of the components of the cooling and fire protection system, and the mass of the components of the dust removal system;

[0106] If the engine does not have a reducer, the mass of the power unit is:

[0107] M cy =n eng K cy M eng

[0108] Where M cy Indicates the mass of the helicopter's power unit if the engine is not equipped with a reducer; n eng Indicates the number of engines installed on the helicopter; M eng Indicates the mass of a single engine; K cy It indicates the engine mass increase factor caused by the conventional installation system components, cooling and fire protection system components and dust removal system components of the gas turbine engine on the helicopter, which can be 1.65 to 1.85;

[0109] If the engine is equipped with a reducer and the influence of the reducer mass needs to be eliminated, the mass of the power unit is:

[0110] M′cy =n eng (K cy M' eng +M red +M sys )

[0111]

[0112] In the formula, M′ cy Indicates the weight of the helicopter's power plant when the engine is equipped with a reducer; M red Indicates the mass of the engine reducer; PW e represents the design input power of the reducer; n2 represents the design input speed of the reducer; M sys The mass of the power unit accessories and system (the mass remains unchanged when the engine mass changes) can be taken as (0.18~0.25)M′ eng ; M′ eng Indicates the engine mass without reduction gear and accessory system.

[0113] Step 3: Based on the basic data of the helicopter-engine system, simulate the helicopter flight trajectory to obtain the helicopter flight fuel quality; including:

[0114] Based on the basic data of the helicopter-engine system, the helicopter flight trajectory is divided into a plurality of flight trajectory sections according to the helicopter flight mission profile given in the design task of the use condition data; wherein the flight conditions and required power of each flight trajectory section of the helicopter remain unchanged;

[0115] The helicopter aerodynamic performance calculation mathematical model is used to calculate the required power of each flight trajectory section of the helicopter; the engine aerodynamic thermodynamic calculation mathematical model is used to calculate the engine fuel consumption rate corresponding to the required power;

[0116] Obtaining the helicopter flight fuel mass according to the required power and its corresponding engine fuel consumption rate;

[0117] The helicopter flight fuel quality (applicable to all helicopters) is:

[0118]

[0119] Among them, M mL represents the helicopter flight fuel mass applicable to all helicopters, t n Indicates the total flight time of the helicopter to complete the flight; M m represents the mass of fuel consumed; r represents the total number of flight trajectory segments; i represents the number of flight trajectory segments; Δt i PW represents the flight time on the i-th flight track segment; iWith sfc i They represent the required power and fuel consumption rate for the i-th flight segment respectively;

[0120] If the helicopter flight trajectory is a simple take-off-cruise at a certain altitude-landing type continuous cruise flight segment, the helicopter flight fuel mass is:

[0121] M' mL =n eng k L sfc P PW P L L / V L

[0122] Among them, M′ mL Indicates the helicopter flight fuel mass when the helicopter flight trajectory is the continuous cruising flight segment; n eng Indicates the number of engines installed on the helicopter; PW p With sfc p They represent the average power and fuel consumption rate of the helicopter during the cruising flight phase; k L The coefficient that takes into account the influence of fuel flow rate in other flight stages other than the cruise stage (such as Figure 3 shown); L L Indicates the helicopter range; V L Indicates the helicopter flight technical speed, which is the average speed considering all flight phases, based on the cruising speed V p Value determination: V L =k V V p , where the correction factor k V according to Figure 3 Select.

[0123] The above is the case when the helicopter's fuel tank and fuel supply system have been confirmed, and only the fuel mass consumed by the helicopter during flight is considered; if the helicopter fuel system device needs to be redesigned and adjusted, it can be redesigned according to the requirements of the helicopter's fuel consumption and engine mass during flight, and the corresponding relationship can be introduced for consideration to maximize the benefits brought by the cycle parameter optimization.

[0124] Step 4: Using the helicopter's power unit mass and the helicopter's flight fuel mass, obtain the helicopter's takeoff mass and perform a test. After the test is passed, calculate the helicopter's efficiency criterion; specifically, the following contents are included:

[0125] Before obtaining the helicopter take-off mass by using the power plant mass and the helicopter flight fuel mass, a first tolerance coefficient ε is used. m Determine the helicopter flight fuel mass M mL or M′ mLWhether the initial given flight fuel mass M is met mLs :

[0126] When the calculated flight fuel mass of the helicopter meets the initially given flight fuel mass, the helicopter take-off mass is calculated and obtained by using the helicopter power unit mass and the helicopter flight fuel mass;

[0127] When the calculated flight fuel mass of the helicopter does not meet the initially given flight fuel mass, returning to step 1, reselecting the value of the engine cycle parameter to be optimized within the initially given range of the optimized engine cycle parameter, and recalculating the helicopter power unit mass and the helicopter flight fuel mass until the helicopter flight fuel mass meets the initially given flight fuel mass;

[0128] Among them, if or When , the calculated helicopter flight fuel mass meets the initial given flight fuel mass; otherwise, it does not meet; where ε m The value can be 0.9 to 1.1.

[0129] The helicopter take-off mass is:

[0130] M oc =M load +M ce +M enp +M me

[0131] Among them, M oc represents the take-off mass of the helicopter; M load Represents the effective load of the helicopter, which is given when the helicopter is designed; M enp Indicates the mass of the helicopter excluding the power unit and fuel system, including the mass of the helicopter body, equipment and crew; M ce The mass M of the helicopter power unit when the engine is not equipped with a reducer cy Or the helicopter power unit mass M' when the engine is equipped with a reducer cy ;M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or the helicopter flight fuel mass M′ when the helicopter flight trajectory is the continuous cruise flight segment mL ;

[0132] Using the second tolerance factor ε m0 , determine the helicopter take-off mass M oc Whether the initial take-off mass M0 is met:

[0133] If the helicopter take-off mass M ocIf the initial given take-off mass M0 is satisfied, the helicopter efficiency criterion is calculated;

[0134] If the helicopter take-off mass M oc If the initial given takeoff mass M0 is not satisfied, then return to step 1, reselect the value of the engine cycle parameter to be optimized within the initial given range of the optimized engine cycle parameter, and calculate the helicopter power unit mass and the helicopter flight fuel mass again until the helicopter takeoff mass M0 is satisfied. oc Meet the initial given take-off mass M0;

[0135] Among them, when When the helicopter take-off mass M oc Satisfies the initial given takeoff mass M0; otherwise, it does not satisfy; where ε m0 The value can be 0.95~1.05.

[0136] Step 5: Determine the optimization optional domain of the engine cycle parameters to be optimized according to the calculation results of the helicopter efficiency criterion; specifically, the following contents are included:

[0137] Helicopter efficiency criteria include: power plant efficiency criteria, transport capacity criteria, transport efficiency criteria and fuel consumption rate criteria;

[0138] Helicopter efficiency criteria also include: helicopter take-off mass or payload, range, helicopter life cycle cost, hanging and installation work productivity, agricultural work productivity, agricultural helicopter economic efficiency and other criteria.

[0139] wherein, according to the mass of the power plant and the mass of the helicopter flight fuel, the efficiency criterion of the power plant is obtained;

[0140] The power plant efficiency criterion is:

[0141]

[0142] In the formula, γ e Indicates the power plant efficiency criterion; M ce The mass M of the helicopter power unit when the engine is not equipped with a reducer cy Or the helicopter power unit mass M' when the engine is equipped with a reducer cy ;M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or when the helicopter flight trajectory is the continuous cruise flight segment, the helicopter flight fuel mass M′ mL ;n eng Indicates the number of engines installed on the helicopter; PW indicates the power of a single engine; K cy Represents the term related to the engine power-to-weight ratio; Ksfc Indicates items related to engine fuel consumption rate;

[0143] The transport capacity criteria are:

[0144]

[0145] In the formula, γ load Indicates the transport capacity criterion; M load is the helicopter's payload; M emp Indicates the unladen mass of the helicopter; V L It represents the flight technical speed, which is the average speed considering all flight phases, according to the cruising speed V p Value determination: V L =k V V p , k V represents the correction factor;

[0146] The transport efficiency criterion is:

[0147]

[0148] In the formula, γ leff represents the transport efficiency criterion; k L V represents the coefficient for considering the influence of fuel flow rate in other flight stages other than the cruise stage; P Indicates cruising speed; M cn Indicates the helicopter crew mass; L L Indicates helicopter range; M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or the helicopter flight fuel mass M' when the helicopter flight trajectory is the continuous cruise flight segment mL ;M emp Indicates the unladen mass of the helicopter.

[0149] The fuel consumption rate criteria are:

[0150]

[0151] In the formula, γ ml Indicates fuel consumption rate criterion; M load is the helicopter's payload; L L Indicates helicopter range; M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or when the helicopter flight trajectory is the continuous cruise flight segment, the helicopter flight fuel mass M′ mL ; ε represents the effective load utilization.

[0152] According to the purpose and flight status characteristics of the helicopter, one or more criteria are comprehensively selected for calculation to obtain corresponding calculated values; according to the calculated values ​​of one or more criteria in the selected helicopter efficiency criteria, the extreme value and compatible range of each criterion are correspondingly determined;

[0153] The calculated values ​​based on one or more criteria, as well as the extreme values ​​and compatible ranges corresponding to each criterion, will satisfy The value range of the engine cycle parameter to be optimized is used as the feasible domain of parameter optimization of the helicopter efficiency criterion; wherein, Y i , Y opt,i , ΔY i Respectively represent the calculated value, extreme value and compatible range of the i-th criterion; the compatible range ΔY of the helicopter efficiency criterion i The applicable range of can be 0.5% to 5%;

[0154] According to the feasible domain of parameter optimization of the helicopter efficiency criterion, and taking into account the engine performance indicators, design technology level, material selection, turbine blade cooling and other condition constraints, the engine cycle parameters to be optimized (compressor pressure ratio π k and the combustion chamber outlet temperature T t4 )'s optimized optional fields.

[0155] Based on the same inventive concept, an embodiment of the present invention also provides a cycle parameter optimization system for an aviation gas turbine shaft engine, such as Figure 4 As shown, it includes a selection unit, an acquisition unit, a calculation unit and a determination unit,

[0156] The selection unit is used to select the engine cycle parameters to be optimized and their corresponding ranges based on the helicopter-engine system basic data;

[0157] The acquisition unit is used to calculate and acquire the mass of the helicopter's power plant according to the helicopter-engine system basic data and the engine cycle parameters to be optimized; and is also used to simulate the helicopter's flight trajectory based on the helicopter-engine system basic data to acquire the helicopter's flight fuel mass;

[0158] The calculation unit is used to obtain the helicopter take-off mass by using the helicopter's power plant mass and flight fuel mass, and to check the helicopter's take-off mass, and to calculate the helicopter efficiency criterion after the check is passed;

[0159] The determination unit is used to determine the optimization optional domain of the engine cycle parameter to be optimized according to the calculation result of the helicopter efficiency criterion.

[0160] Regarding the system in the above embodiment, the specific manner in which each unit module performs operations has been described in detail in the embodiment of the method, and will not be elaborated here.

[0161] Based on the same inventive concept, an embodiment of the present invention further provides an electronic device, whose structure is as follows: Figure 5 As shown, it includes: a memory, a processor, and a computer program stored in the memory and running on the processor, and when the processor executes the computer program, it implements the aforementioned method for optimizing the cycle parameters of an aviation gas turbine shaft engine.

[0162] Based on the same inventive concept, an embodiment of the present invention further provides a computer storage medium, wherein the computer storage medium stores computer executable instructions, and when the computer executable instructions are executed, the aforementioned method for optimizing cycle parameters of an aviation gas turbine shaft engine is implemented.

[0163] Finally, it should be noted that the above is only a preferred embodiment of the present invention and is not intended to limit the present invention. Although the present invention has been described in detail with reference to the aforementioned embodiments, it is still possible for those skilled in the art to modify the technical solutions described in the aforementioned embodiments or to make equivalent substitutions for some of the technical features therein. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A method for optimizing cycle parameters of an aviation gas turbine shaft engine, It is characterized in that include: Based on the basic data of the helicopter-engine system, select the engine cycle parameters to be optimized and their corresponding ranges; Determining the power unit mass of the helicopter based on the helicopter-engine system basic data and the engine cycle parameters to be optimized; Based on the basic data of the helicopter-engine system, simulate the flight trajectory of the helicopter to obtain the flight fuel quality of the helicopter; Using the helicopter's power unit mass and flight fuel mass, obtain the helicopter's takeoff mass and conduct a test. After the test is passed, calculate the helicopter's efficiency criterion; According to the calculation results of the helicopter efficiency criterion, the optimization optional domain of the engine cycle parameters to be optimized is determined; Wherein, determining the power unit mass of the helicopter according to the helicopter-engine system basic data and the engine cycle parameters to be optimized includes: Obtaining flow channel dimensions of engine components according to the helicopter-engine system basic data and the engine cycle parameters to be optimized; Determine the engine mass based on the relationship between the flow channel size of the engine components and the mass of the corresponding components; According to the engine mass, obtaining the power unit mass; The method of simulating the helicopter flight trajectory based on the helicopter-engine system basic data and obtaining the helicopter flight fuel quality includes: Based on the helicopter-engine system basic data, dividing the helicopter flight trajectory into a plurality of flight trajectory segments according to the helicopter flight mission profile; The helicopter aerodynamic performance calculation mathematical model is used to calculate the required power for each section of the helicopter flight trajectory; Calculate the engine fuel consumption rate corresponding to the required power by using the engine aerodynamic thermodynamic calculation mathematical model; Obtaining the helicopter flight fuel mass according to the required power and its corresponding engine fuel consumption rate; The step of determining the optimization optional domain of the engine cycle parameter to be optimized according to the calculation result of the helicopter efficiency criterion includes: According to the purpose and flight status characteristics of the helicopter, one or more criteria are selected for calculation to obtain corresponding calculation values; According to the calculated values ​​of one or more criteria in the selected helicopter efficiency criteria, the extreme value and compatible range of each criterion are correspondingly determined; The calculated values ​​based on one or more criteria, as well as the extreme values ​​and compatible ranges corresponding to each criterion, will satisfy The value range of the engine cycle parameter to be optimized is used as the feasible domain of parameter optimization of the helicopter efficiency criterion; wherein, Y i , Y opt,i , ΔY i They represent the calculated value, extreme value and compatible range of the i-th criterion respectively; Determining an optimization optional domain of the engine cycle parameters to be optimized according to the parameter optimization feasible domain of the helicopter efficiency criterion; The helicopter efficiency criteria include a power plant efficiency criterion, a transport capacity criterion, a transport efficiency criterion and a fuel consumption rate criterion; The power plant efficiency criterion is obtained according to the power plant mass and the helicopter flight fuel mass.

2. A method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 1, It is characterized in that The engine cycle parameters to be optimized are selected based on the helicopter-engine system basic data, including: The helicopter-engine system basic data includes: engine basic data, helicopter basic data, use condition data and engine cycle parameters to be optimized; The engine cycle parameters to be optimized include compressor pressure ratio and combustion chamber outlet temperature.

3. The method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 1, It is characterized in that The step of obtaining the flow channel size of the engine component according to the helicopter-engine system basic data and the engine cycle parameters to be optimized includes: According to the basic data of the helicopter-engine system and the engine cycle parameters to be optimized, the inlet and outlet cross-sectional parameters and performance parameters of each engine component are calculated using the engine aerodynamic thermodynamic calculation mathematical model; Determine the flow channel dimensions of engine components based on the inlet and outlet cross-sectional parameters and performance parameters of each engine component; Among them, the engine components include: air intake, compressor, gas turbine, combustion chamber, power turbine, casing and transition section.

4. The method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 1, It is characterized in that The step of obtaining the power unit mass according to the engine mass includes: Determining the mass of the power unit based on the mass of the engine, the mass of the components of the conventional installation system of the gas turbine engine on the helicopter, the mass of the components of the cooling and fire protection system, and the mass of the components of the dust removal system; If the engine does not have a reducer, the mass of the power unit is: M cy =n eng K cy M eng Where M cy Indicates the mass of the helicopter's power unit when the engine is not equipped with a reducer; n eng Indicates the number of engines installed on the helicopter; M eng Indicates the mass of a single engine; K cy It represents the engine mass increase factor caused by considering the conventional installation system components, cooling and fire protection system components and dust removal system components of the gas turbine engine on the helicopter; If the engine is equipped with a reducer, the mass of the power unit is: M c ' y =n eng (K cy M e ' ng +M red +M sys ) In the formula, M′ cy Indicates the weight of the helicopter's power plant when the engine is equipped with a reducer; M red Indicates the mass of the engine reducer; PW e Indicates the design input power of the reducer; n 2 Indicates the reducer input design speed; M sys Indicates the mass of power unit accessories and systems; M' eng Indicates the engine mass without reduction gear and accessory system.

5. The method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 1, It is characterized in that The helicopter flight fuel mass is: Among them, M mL represents the helicopter flight fuel mass applicable to all helicopters, t n Indicates the total flight time of the helicopter to complete the flight; M m represents the mass of fuel consumed; r represents the total number of flight trajectory segments; i represents the number of flight trajectory segments; Δt i PW represents the flight time on the i-th flight track segment; i With sfc i They represent the required power and fuel consumption rate for the i-th flight segment respectively; If the helicopter flight trajectory is a simple take-off-cruise at a certain altitude-landing type continuous cruise flight segment, the helicopter flight fuel mass is: M' mL =n eng k L sfc P PW P L L / V L Among them, M′ mL Indicates the helicopter flight fuel mass when the helicopter flight trajectory is the continuous cruising flight segment; n eng Indicates the number of engines installed on the helicopter; PW p With sfc p They represent the average power and fuel consumption rate of the helicopter during the cruising flight phase; k L It represents the coefficient that considers the influence of fuel flow rate in other flight stages other than the cruise stage; L L Indicates the helicopter range; V L Indicates the helicopter flight technical speed, which is the average speed considering all flight phases, according to the cruising speed V p Value determination: V L =k V V p , k V Indicates the correction factor.

6. A method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 1 or 5, It is characterized in that Before obtaining the helicopter take-off mass by using the power plant mass and the helicopter flight fuel mass, a first tolerance coefficient ε is used. m Determine the helicopter flight fuel mass M mL or M′ mL Whether the initial given flight fuel mass M is met mLs : When the calculated flight fuel mass of the helicopter meets the initially given flight fuel mass, the helicopter take-off mass is calculated and obtained by using the helicopter power unit mass and the helicopter flight fuel mass; When the calculated flight fuel mass of the helicopter does not meet the initially given flight fuel mass, reselecting the value of the engine cycle parameter to be optimized within the initially given range of the optimized engine cycle parameter, and recalculating the helicopter power unit mass and the helicopter flight fuel mass until the helicopter flight fuel mass meets the initially given flight fuel mass; Among them, if or , the calculated flight fuel mass of the helicopter meets the initially given flight fuel mass.

7. A method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 6, It is characterized in that The helicopter take-off mass is obtained and tested using the helicopter power unit mass and the helicopter flight fuel mass. After the test is passed, the helicopter efficiency criteria are calculated, including: The helicopter take-off mass is: M oc =M load +M ce +M enp +M me Among them, M oc represents the take-off mass of the helicopter; M load Represents the effective load of the helicopter, which is given when the helicopter is designed; M enp Indicates the mass of the helicopter excluding the power unit and fuel system, including the mass of the helicopter body, equipment and crew; M ce The mass M of the helicopter power unit when the engine is not equipped with a reducer cy Or the helicopter power unit mass M' when the engine is equipped with a reducer cy ;M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or the helicopter flight fuel mass M′ when the helicopter flight trajectory is the continuous cruise flight segment mL ; Using the second tolerance factor ε m0 , determine the helicopter take-off mass M oc Does it meet the initial given takeoff mass M? 0 : If the helicopter take-off mass M oc Satisfy the initial given takeoff mass M 0 , then calculate the helicopter efficiency criterion; If the helicopter take-off mass M oc Does not meet the initial given takeoff mass M 0 , then within the initially given range of the optimized engine cycle parameters, the values ​​of the engine cycle parameters to be optimized are reselected, and the helicopter power unit mass and the helicopter flight fuel mass are calculated again until the helicopter takeoff mass M oc Satisfy the initial given takeoff mass M 0 ; Among them, when When the helicopter take-off mass M oc Satisfy the initial given takeoff mass M 0 .

8. The method for optimizing cycle parameters of an aviation gas turbine shaft engine according to claim 1, It is characterized in that The power plant efficiency criterion is: Among them, γ e Indicates the power plant efficiency criterion; M ce The mass M of the helicopter power unit when the engine is not equipped with a reducer cy Or the helicopter power unit mass M' when the engine is equipped with a reducer cy ;M me Denotes the helicopter flight fuel mass M applicable to all helicopters mL Or when the helicopter flight trajectory is the continuous cruise flight segment, the helicopter flight fuel mass M′ mL ;n eng Indicates the number of engines installed on the helicopter; PW indicates the power of a single engine; K cy Represents the term related to the engine power-to-weight ratio; K sfc Indicates items related to engine fuel consumption rate; The transport capacity criteria are: Among them, γ load Indicates the transport capacity criterion; M load Indicates the helicopter's payload; M emp Indicates the unladen mass of the helicopter; V L It represents the flight technical speed, which is the average speed considering all flight phases, according to the cruising speed V p Value determination: V L =k V V p , k V represents the correction factor; The transport efficiency criterion is: Among them, γ leff represents the transport efficiency criterion; k L V represents the coefficient for considering the influence of fuel flow rate in other flight stages other than the cruise stage; P Indicates cruising speed; M cn Indicates the helicopter crew mass; L L Indicates helicopter range; M emp Indicates the unladen mass of the helicopter; The fuel consumption rate criteria are: Among them, γ ml represents the fuel consumption rate criterion; ε represents the effective load utilization rate.

9. A cycle parameter optimization system for an aviation gas turbine shaft engine, It is characterized in that It includes a selection unit, an acquisition unit, a calculation unit and a determination unit. The selection unit is used to select the engine cycle parameters to be optimized and their corresponding ranges based on the helicopter-engine system basic data; The acquisition unit is used to calculate and acquire the mass of the helicopter's power plant according to the helicopter-engine system basic data and the engine cycle parameters to be optimized; and is also used to simulate the helicopter's flight trajectory based on the helicopter-engine system basic data to acquire the helicopter's flight fuel mass; The calculation unit is used to obtain the helicopter take-off mass by using the helicopter's power plant mass and flight fuel mass, and to check the helicopter's take-off mass, and to calculate the helicopter efficiency criterion after the check is passed; The determination unit is used to determine the optimization optional domain of the engine cycle parameter to be optimized according to the calculation result of the helicopter efficiency criterion; Wherein, determining the power unit mass of the helicopter according to the helicopter-engine system basic data and the engine cycle parameters to be optimized includes: Obtaining flow channel dimensions of engine components according to the helicopter-engine system basic data and the engine cycle parameters to be optimized; Determine the engine mass based on the relationship between the flow channel size of the engine components and the mass of the corresponding components; According to the engine mass, obtaining the power unit mass; The method of simulating the helicopter flight trajectory based on the helicopter-engine system basic data and obtaining the helicopter flight fuel quality includes: Based on the helicopter-engine system basic data, dividing the helicopter flight trajectory into a plurality of flight trajectory segments according to the helicopter flight mission profile; The helicopter aerodynamic performance calculation mathematical model is used to calculate the required power for each section of the helicopter flight trajectory; Calculate the engine fuel consumption rate corresponding to the required power by using the engine aerodynamic thermodynamic calculation mathematical model; Obtaining the helicopter flight fuel mass according to the required power and its corresponding engine fuel consumption rate; The step of determining the optimization optional domain of the engine cycle parameter to be optimized according to the calculation result of the helicopter efficiency criterion includes: According to the purpose and flight status characteristics of the helicopter, one or more criteria are selected for calculation to obtain corresponding calculation values; According to the calculated values ​​of one or more criteria in the selected helicopter efficiency criteria, the extreme value and compatible range of each criterion are correspondingly determined; The calculated values ​​based on one or more criteria, as well as the extreme values ​​and compatible ranges corresponding to each criterion, will satisfy The value range of the engine cycle parameter to be optimized is used as the feasible domain of parameter optimization of the helicopter efficiency criterion; wherein, Y i , Y opt,i , ΔY i They represent the calculated value, extreme value and compatible range of the i-th criterion respectively; Determining an optimization optional domain of the engine cycle parameters to be optimized according to the parameter optimization feasible domain of the helicopter efficiency criterion; The helicopter efficiency criteria include a power plant efficiency criterion, a transport capacity criterion, a transport efficiency criterion and a fuel consumption rate criterion; The power plant efficiency criterion is obtained according to the power plant mass and the helicopter flight fuel mass.

10. An electronic device, It is characterized in that include: A memory, a processor, and a computer program stored in the memory and running on the processor, wherein when the processor executes the computer program, a method for optimizing cycle parameters of an aviation gas turbine shaft engine as described in any one of claims 1-8 is implemented.

11. A computer storage medium, It is characterized in that The computer storage medium stores computer executable instructions, which, when executed, implement a method for optimizing cycle parameters of an aviation gas turbine shaft engine as described in any one of claims 1-8.

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