Method for calculating installation loss of turboshaft engine

Abstract

The invention relates to the technical field of engines, and provides a method for calculating installation loss of a turboshaft engine, which comprises the following steps of: acquiring second input data from a turboshaft engine performance calculation model established on the basis of bench data, acquiring first input data, acquiring first sub-input data on the basis of first total temperature and first total pressure, and acquiring second sub-input data on the basis of second total pressure; the first sub-input data comprises a target gas turbine outlet total temperature and target gas turbine outlet total pressure, and the first sub-input data and the first input data are the same in calculation mode; if the conditions are met, the target gas turbine outlet total temperature is equal to the second total temperature, the target gas turbine outlet total pressure is equal to the second total pressure, the second flow is equal to the first flow, the power of the power turbine is equal to the power required by the helicopter, and the parallel operation rule condition is met; and obtaining the power of the power turbine and the required power of the helicopter under the non-installation condition to obtain the corresponding installation loss. The installation loss obtained through the method has the capability of evaluating and optimizing the needed installation loss in the design stage.

Description

Technical Field

[0001] The present invention belongs to the technical field of engines, and particularly relates to a method for calculating the installation loss of a turboshaft engine. Background Art

[0002] The system composed of a helicopter and an engine is a complex system, and its structure, performance, flow, etc. are closely coupled with each other. When a turboshaft engine is installed on a helicopter, under the same compressor speed or gas turbine outlet temperature conditions, the installed power of the turboshaft engine is usually less than the bench power of the turboshaft engine (the power of the turboshaft engine installed under the test bench. Compared with being installed on a helicopter, installing on the test bench makes the intake and exhaust of the turboshaft engine free from the interference of external air flow), that is, the installation loss of the turboshaft engine is generated. Moreover, the installation losses of different types of helicopters can vary from a few percent to more than ten percent. Excessive installation losses have an extremely adverse impact on flight performance. How to avoid possible increased installation losses and improve flight performance during the design and use stages requires research on the installation loss simulation method in order to conduct improvement research on reducing the installation loss.

[0003] According to the publicly available information, for the assessment of the installation loss of a turboshaft engine, it is basically based on the conversion and comparison of bench data and flight data. The assessment based on the geometric model during the design stage is still blank; the publicly disclosed internal and external flow simulation methods and processes of a helicopter / turboshaft engine can only provide partial reference for the installation loss, but the calculation method of the installation loss cannot be inferred from this method. In addition, the installation loss calculation method provided by Patent CN108061659B is completely a processing method based on test data, which belongs to a posteriori method and does not have the ability of assessment and optimization required during the design stage. At the same time, the power normalization method of the turboshaft engine adopted is inappropriate because the similarity conversion of the turboshaft engine power does not satisfy the similarity conversion rules. Summary of the Invention

[0004] In view of the above problems, the present invention proposes a method for calculating the installation loss of a turboshaft engine, and the method includes:

[0005] Obtaining second input data from a performance calculation model of a turboshaft engine established based on bench data, where the second input data includes the first flow rate at the inlet of the turboshaft engine section, the first total temperature and the first total pressure at the inlet of the gas turbine;

[0006] Obtaining first input data, where the first input data includes the second total temperature, the second total pressure and the second flow rate of the turboshaft engine section, as well as the power of the power turbine and the power required by the helicopter;

[0007] The first sub-input data is obtained based on the first total temperature and the first total pressure. The first sub-input data includes the total temperature at the outlet of the target gas turbine and the total pressure at the outlet of the target gas turbine, and the first sub-input data is obtained by the same calculation method as the first input data;

[0008] If the following conditions are met: the total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power of the power turbine is equal to the power required by the helicopter, and the conditions of the parallel operation rule are satisfied, then the corresponding installation loss is obtained according to the power of the power turbine under non-installed conditions and the power required by the helicopter.

[0009] Optionally, before obtaining the second input data from the performance calculation model of the turboshaft engine established based on the test bench data, it includes:

[0010] Select the numerical value of the dependent variable;

[0011] According to the numerical value of the dependent variable and the performance calculation method of the turboshaft engine, a performance calculation model of the turboshaft engine is established based on the test bench data.

[0012] Optionally, obtaining the second input data from the performance calculation model of the turboshaft engine established based on the test bench data includes:

[0013] Obtain the calculation input conditions from the performance calculation model of the turboshaft engine established based on the test bench data;

[0014] Obtain the second input data according to the calculation input conditions.

[0015] Optionally, obtaining the first input data includes:

[0016] Obtain the outlet boundary conditions and the inlet boundary conditions;

[0017] Calculate the first input data according to the outlet boundary conditions and the inlet boundary conditions.

[0018] Optionally, obtaining the outlet boundary conditions and the inlet boundary conditions includes:

[0019] Define the cross-section of the inlet duct outlet / compressor inlet as the cross-section of the turboshaft engine, and the corresponding first data required;

[0020] Define the cross-section of the gas turbine outlet / power turbine inlet as the cross-section of the turboshaft engine, and the corresponding second data required;

[0021] Set the outlet boundary conditions and the inlet boundary conditions based on the first data and the second data.

[0022] Optionally, the method further includes:

[0023] If the following conditions are not met:

[0024] The total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power of the power turbine is equal to the power required by the helicopter, and step two is re-executed until all conditions are met.

[0025] Optionally, the non-satisfaction conditions include:

[0026] If it does not satisfy that the total temperature at the outlet of the target gas turbine is equal to the second total temperature and the total pressure at the outlet of the target gas turbine is equal to the second total pressure, then adjust the total temperature and total pressure in the first input data until the conditions are met;

[0027] If it does not satisfy:

[0028] The second flow rate is equal to the first flow rate, then adjust the static pressure in the first input data until the conditions are met;

[0029] If it does not satisfy:

[0030] The power of the power turbine is equal to the power required by the helicopter, then adjust the compressor speed until the conditions are met.

[0031] Optionally, if the parallel operation rule conditions are met, including:

[0032] The gas generator speeds of the left and right corresponding turboshaft engines are equal, or the power turbine torques of the left and right corresponding turboshaft engines are equal, otherwise execute step two until the conditions are met.

[0033] Optionally, if the conditions are met: the total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, and the power of the power turbine is equal to the power required by the helicopter, then obtain the corresponding installation loss according to the power of the power turbine and the power required by the helicopter under non-installed conditions, including:

[0034] Calculate the installation loss at constant speed and the installation loss at constant temperature.

[0035] The present invention provides a device for calculating the installation loss of a turboshaft engine, including:

[0036] An acquisition module for acquiring second input data from a turboshaft engine performance calculation model established based on bench data, where the second input data includes the first flow rate at the inlet of the turboshaft engine section and the first total temperature and first total pressure at the inlet of the gas turbine;

[0037] A first calculation module for acquiring first input data, where the first input data includes the second total temperature, second total pressure and second flow rate of the turboshaft engine section, as well as the power of the power turbine and the power required by the helicopter;

[0038] A second calculation module, configured to obtain first sub-input data based on the first total temperature and the first total pressure, where the first sub-input data includes the total temperature at the outlet of the target gas turbine and the total pressure at the outlet of the target gas turbine, and the first sub-input data is obtained by the same calculation method as the first input data;

[0039] A condition module, configured to, if the following conditions are met: the total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power of the power turbine is equal to the power required by the helicopter, and the parallel operation rule conditions are met, obtain the corresponding installation loss according to the power of the power turbine under non-installed conditions and the power required by the helicopter.

[0040] The method for calculating the installation loss of a turboshaft engine provided by the present invention has the following advantages compared with the prior art:

[0041] By constructing a turboshaft engine performance calculation model and a turboshaft engine parallel operation constraint, performance parameters such as the installation power, total temperature at the outlet of the gas turbine, and total pressure of the turboshaft engine are obtained; the first sub-input data is obtained by the same calculation method as the first input data, that is, the total temperature and total pressure at the outlet of the gas turbine are used as input conditions and substituted into the corresponding calculation method. Finally, after meeting the corresponding conditions, the power of the power turbine under non-installed conditions can be obtained, and the corresponding installation loss can be obtained according to the power of the power turbine and the power required by the helicopter. The installation loss obtained by this method has the ability to evaluate and optimize the required installation loss during the design stage.

[0042] Other features and advantages of the present invention will be described in the following specification, and, in part, will be obvious from the specification, or will be understood by implementing the present invention. The objectives and other advantages of the present invention can be achieved and obtained by the structure pointed out in the specification and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0044] Figure 1 shows a schematic flow chart of the method for calculating the installation loss of a turboshaft engine in an embodiment of the present invention;

[0045] Figure 2 shows a schematic cross-sectional view of the device for calculating the installation loss of a turboshaft engine in an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0046] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are 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.

[0047] like Figure 1 As shown, the present invention provides a method for calculating the installation loss of a turboshaft engine, the method comprising:

[0048] Step S10, obtaining second input data from a turboshaft engine performance calculation model established based on bench data, the second input data including a first flow rate at the turboshaft engine cross section inlet and a first total temperature and a first total pressure at the gas turbine inlet. It should be noted that the present invention is a twin-turbo engine, therefore, the first flow rate includes two turboshaft engine cross section inlet flows on the left and right sides of the engine body, and the position of the first flow rate is different from the position where the first total temperature and the first total pressure at the gas turbine inlet are generated, for specific reference Figure 2 The first flow position at the inlet of the turboshaft engine section is indicated by the label 2, and the first total temperature and the first total pressure position at the inlet of the gas turbine section is indicated by the label 45, that is, two different positions.

[0049] In one embodiment, before obtaining the second input data from establishing a turboshaft engine performance calculation model based on the bench data, the method includes:

[0050] Select the value of the dependent variable;

[0051] According to the value of the dependent variable and the turboshaft engine performance calculation method, a turboshaft engine performance calculation model based on the bench data is established. Specifically, the dependent variable is selected as follows Figure 2 The compressor outlet to the gas turbine first-stage guide vane bleed air coefficient x1 (10% in the figure), the compressor outlet to the gas turbine second-stage guide vane bleed air coefficient x2 (4% in the figure), the compressor intermediate stage to the power turbine first-stage guide vane bleed air coefficient x3 and the compressor intermediate stage to the power turbine second-stage guide vane bleed air coefficient x4, the compressor component flow characteristic correction factor x5, the efficiency characteristic correction factor x6, the gas turbine component flow characteristic correction factor x7, the efficiency characteristic correction factor x8, the power turbine component flow characteristic correction factor x9, the efficiency characteristic correction factor x 10 The target value variables are the bench data of the turboshaft engine and the calculated performance value of the turboshaft engine (including the fuel flow rate wfb and the physical speed of the power turbine n p As input variable (state-determining parameter), the gas turbine speed n g, compressor pressure ratio π, compressor inlet flow rate Wa2, power turbine output power PW, power turbine inlet temperature T t45 , tailpipe inlet temperature T t6 ) weighted quantity of relative error, using a genetic algorithm to optimize the dependent variable to minimize the target value variable, thereby obtaining the value of the dependent variable, which is combined with the traditional performance calculation method of a turboshaft engine to form the performance calculation method of a turboshaft engine based on bench test data, that is, the performance calculation model of a turboshaft engine.

[0052] In one embodiment, obtaining second input data from the performance calculation model of a turboshaft engine established based on bench test data, including:

[0053] Obtaining the calculation input conditions from the performance calculation model of a turboshaft engine established based on bench test data. By extracting the gas generator calculation model from the performance calculation model of a turboshaft engine established based on bench test data, the calculation input conditions can be obtained from the gas generator calculation model, where the calculation input conditions include the total temperature T at the inlet of the intake duct / compressor inlet section t2,L and T t2,R , total pressure P t2,L and P t2,R , outlet condition flow rate Wa 45,L and Wa 45,R . Wherein, L and R represent the left and right engines respectively.

[0054] Obtaining the second input data according to the calculation input conditions. Calculating the second input data according to the above parameters, the specific calculation process and corresponding calculation method are well-known in the industry and will not be elaborated. Finally, the second input data is obtained, and the inlet flow rate (i.e., the first flow rate) Wa of section 2 of the turboshaft engine is obtained 2,L,eng and Wa 2,L,eng and the total temperature (the first total temperature) T at section 45 of the gas turbine inlet t45,L,eng and T t45,R,eng , total pressure (the first total pressure) P t45,L,eng and P t45,R,eng .

[0055] Step S20, obtaining the first input data, the first input data including the second total temperature, second total pressure and second flow rate of the turboshaft engine section, as well as the power turbine power and the helicopter demand power. It should be noted that the first input data is obtained by calculating the computational domain A using three-dimensional CFD (the corresponding calculation method is well-known in the industry and will not be elaborated). The second total temperature, second total pressure and second flow rate of the turboshaft engine section respectively correspond to the total temperature T at section 2 of the turboshaft engine t2,L and T t2,R , total pressure P t2,L and P t2,R , flow rate Wa 2,L,CFD and Wa2,R,CFD (Using CFD as a subscript is to indicate that this parameter is the result of CFD calculation), the power of the power turbine and the power required by the helicopter correspond to the power PW of the power turbine at the 45th section of the turboshaft engine L and PW R , the power required by the helicopter PW hel , in addition to the above parameters, it also includes the torque q of the power turbine L and q R .

[0056] In one embodiment, obtaining the first input data includes:

[0057] Obtaining the outlet boundary condition and the inlet boundary condition. The outlet boundary condition is expressed as setting the static pressures P at the outlets of the left and right air inlets / compressor inlets simultaneously s2,L and P s2,R , the inlet boundary condition is expressed as the total temperature T at the outlet of the gas turbine / inlet of the power turbine t45,L and T t45,R , the total pressure P t45,L and P t45,R . All of these parameters are the basic parameters for calculating the first input data. It should be noted that a computational grid is divided for the helicopter, the external control volume, and the left and right turboshaft engine air inlets, power turbines, and tail nozzles to form a computational region. The geometric boundary of this region is set as a pressure far-field boundary. According to the flight state, parameters such as the flight Mach number, atmospheric pressure, and temperature of the oncoming flow (airflow) are specified, and these parameters need to be set in this region or at the pressure far-field boundary.

[0058] The first input data is calculated based on the outlet boundary condition and the inlet boundary condition. It should be noted that this calculation process is well-known.

[0059] In one embodiment, obtaining the outlet boundary condition and the inlet boundary condition includes:

[0060] Define the outlet of the air inlet / compressor inlet section as the section of the turboshaft engine, and the corresponding first data required. It should be noted that first, define the outlet of the air inlet / compressor inlet section as the 2nd section of the turboshaft engine, and its symbols are the total temperature T t2 , the total pressure P t2 , the static pressure P s2 , and the flow rate Wa2.

[0061] Define the outlet of the gas turbine / inlet of the power turbine as the section of the turboshaft engine, and the corresponding second data required. It should be noted that define the outlet of the gas turbine / inlet of the power turbine as the 45th section of the turboshaft engine, and its symbols are the total temperature T t45 , the total pressure P t45 , the flow rate Wa 45 .

[0062] Set the outlet boundary condition and the inlet boundary condition based on the first data and the second data.

[0063] Step S30: Obtain first sub-input data based on the first total temperature and the first total pressure. The first sub-input data includes the target gas turbine outlet total temperature and the target gas turbine outlet total pressure, and the first sub-input data is obtained by the same calculation method as the first input data. It should be noted that in this step, the first total temperature and the first total pressure are input into the formula in which the first input data is calculated, that is, it is equivalent to inputting the first total temperature and the first total pressure into the turboshaft engine performance model corrected based on the test bench performance data for calculation, so that the first sub-input data can be compared with the first data. If the conditions are not met, adjustments are made.

[0064] Step S40: If the following conditions are met: the target gas turbine outlet total temperature is equal to the second total temperature, the target gas turbine outlet total pressure is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power turbine power is equal to the helicopter demand power, and the parallel operation rule conditions are met, then obtain the corresponding installation loss according to the power turbine power under non-installed conditions and the helicopter demand power. By meeting the conditions, the power turbine power under non-installed conditions can be obtained, and the corresponding installation loss can be obtained according to the power turbine power and the helicopter demand power. The installation loss obtained by this method has the ability to evaluate and optimize the required installation loss in the design stage. In addition, since this application is for a twin-engine, therefore, the flow rate, total temperature, total pressure, and power turbine power are all two for the left and right sides. For example, the sum of the two power turbine powers is equal to the helicopter demand power, or the sum of the two flow rates in the second flow rate is equal to the sum of the two flow rates in the first flow rate.

[0065] In one embodiment, the method further includes:

[0066] If the following conditions are not met:

[0067] The target gas turbine outlet total temperature is equal to the second total temperature, the target gas turbine outlet total pressure is equal to the second total pressure, the second flow rate is equal to the first flow rate, and the power turbine power is equal to the helicopter demand power, then re-execute step two until all conditions are met.

[0068] In one embodiment, not meeting the conditions includes:

[0069] Not meeting: the target gas turbine outlet total temperature is equal to the second total temperature, the target gas turbine outlet total pressure is equal to the second total pressure, then adjust the total temperature and total pressure in the first input data until the conditions are met. If these two conditions are not met, that is, the target gas turbine outlet total temperature is equal to the second total temperature, the target gas turbine outlet total pressure is equal to the second total pressure, it is necessary to adjust the total temperature and total pressure at the turbine inlet section of the two engines respectively, that is, adjust the total temperature T in the first input datat45,L With T t45,R , total pressure P t45,L and P t45,R , after the adjustment is completed, re - execute step two until both of these two conditions are met in this step.

[0070] Not satisfied:

[0071] If the second flow rate is equal to the first flow rate, then adjust the static pressure in the first input data and re - execute step two until the condition is met. If the second flow rate is not equal to the first flow rate, then it is necessary to adjust the static pressure P s2,L and P s2,R until the condition is met.

[0072] Not satisfied:

[0073] If the power of the power turbine is equal to the power required by the helicopter, then adjust the compressor speed until the condition is met. It should be noted that if the power of the power turbine is not equal to the power required by the helicopter, it is necessary to adjust the physical speeds n g,L and n g,R , and re - execute step two until the condition is met.

[0074] In one embodiment, the conditions for meeting the parallel - engine rules include:

[0075] The rotational speeds of the gas generators of the left and right corresponding turboshaft engines are equal, or the torques of the power turbines of the left and right corresponding turboshaft engines are equal. Otherwise, execute step two until the condition is met. Optionally, the conditions for meeting the parallel - engine rules also include: the total inlet temperature T t45,L of the left and right power turbines and T t45,R are equal. If any of the above conditions does not meet the parallel - engine rules, then it is necessary to correct the physical speeds n g,L and n g,R until the condition is met. It should be noted that for the parallel - engine rules of turboshaft engines, for helicopters with twin - engines and triple - engines (such as the twin - engine parallel - engine of Mi - 26 and the triple - engine parallel - engine of CH53K), there are 2 - 3 turboshaft engines providing power to the helicopter rotor and tail rotor after parallel - engine on the main reduction gear. Generally, the rules for engine parallel - engine use equal compressor physical speed n g , equal power turbine torque q pt , and equal total temperature T t45 at the gas turbine outlet.

[0076] In one embodiment, if the following conditions are met: the total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, and the power of the power turbine is equal to the power required by the helicopter, then the installation loss corresponding to the power of the power turbine under non-installed conditions and the power required by the helicopter is obtained, including:

[0077] Calculate the installation loss at constant rotational speed and the installation loss at constant temperature. Specifically, satisfying the internal and external flow matching (i.e., flow rate condition), power balance (i.e., the power of the power turbine is equal to the power required by the helicopter), and the parallel operation rule (i.e., whether the physical rotational speeds n g,L and n g,R of the left and right compressors are equal, or whether the power torques q L and q R of the left and right power turbines are equal, or whether the total temperatures T t45,L and T t45,R at the inlets of the left and right power turbines are equal) constraints, take the physical rotational speeds n g,L and n g,R of the left and right compressors, and the flight conditions as input parameters, and calculate the power PW L,uninst and PW R,uninst of the power turbine under non-installed conditions according to the performance calculation model of the turboshaft engine. According to the installation loss formula k = (PW uninst -PW) / PW uninst , calculate the installation losses k L and k R of the left and right turboshaft engines respectively, which is called the installation loss at constant rotational speed; take the total temperatures T t45,L and T t45,R at the inlets of the left and right power turbines that satisfy the internal and external flow matching, power balance, and parallel operation rule constraints, and the flight conditions as input parameters, and calculate the power PW L,uninst and PW R,uninst of the power turbine under non-installed conditions according to the performance calculation model of the turboshaft engine. According to the installation loss formula k = (PW uninst -PW) / PW uninst , calculate the installation losses k L and k R of the left and right turboshaft engines respectively, which is called the installation loss at constant temperature.

[0078] It should be noted that if only one turboshaft engine is installed on the helicopter, the left and right engines in the text should be changed to a single engine, and at the same time, the parallel operation rule constraint in the conditions should be deleted, while keeping the helicopter demand power equal to the power turbine power. If three turboshaft engines are installed on the helicopter, the left and right engines in the text should be changed to the left, middle, and right engines, and at the same time, the condition should be changed to whether the helicopter demand power is equal to the sum of the power turbine powers of the three turboshaft engines, whether the gas generator speeds of the left, middle, and right turboshaft engines are pairwise equal, or whether the total outlet temperatures of the gas generators of the left, middle, and right turboshaft engines are pairwise equal, or whether the power turbine torques of the left, middle, and right turboshaft engines are pairwise equal.

[0079] The present invention also provides a device for calculating the installation loss of a turboshaft engine, including:

[0080] An acquisition module, configured to obtain second input data from a performance calculation model of a turboshaft engine established based on bench data, where the second input data includes the first flow rate at the inlet of the turboshaft engine section, the first total temperature, and the first total pressure at the inlet of the gas turbine.

[0081] A first calculation module, configured to obtain first input data, where the first input data includes the second total temperature, the second total pressure, and the second flow rate at the turboshaft engine section, as well as the power turbine power and the helicopter demand power.

[0082] A second calculation module, configured to obtain first sub-input data based on the first total temperature and the first total pressure, where the first sub-input data includes the target total temperature at the outlet of the gas turbine and the target total pressure at the outlet of the gas turbine, and the first sub-input data is obtained through the same calculation method as the first input data.

[0083] A condition module, configured to, if the conditions are met: the target total temperature at the outlet of the gas turbine is equal to the second total temperature, the target total pressure at the outlet of the gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power turbine power is equal to the helicopter demand power, and the parallel operation rule condition is met, obtain the corresponding installation loss according to the power turbine power and the helicopter demand power under non-installation conditions. Through each module, the corresponding installation loss is obtained according to the power turbine power and the helicopter demand power, so that the obtained installation loss has the ability to evaluate and optimize the required installation loss in the design stage.

[0084] Although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: they can still modify the technical solutions described in the foregoing embodiments, or perform equivalent replacements for some of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for calculating the installation loss of a turboshaft engine, characterized in that The method includes: Obtaining second input data from a performance calculation model of a turboshaft engine established based on bench data, where the second input data includes the first flow rate at the inlet of the turboshaft engine section, the first total temperature and the first total pressure at the inlet of the gas turbine; Obtaining first input data, where the first input data includes the second total temperature, the second total pressure and the second flow rate at the turboshaft engine section, as well as the power of the power turbine and the power required by the helicopter; Obtaining first sub-input data based on the first total temperature and the first total pressure, where the first sub-input data includes the target total temperature at the outlet of the gas turbine and the target total pressure at the outlet of the gas turbine, and the first sub-input data is obtained by the same calculation method as the first input data; If the conditions are met: the target total temperature at the outlet of the gas turbine is equal to the second total temperature, the target total pressure at the outlet of the gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power of the power turbine is equal to the power required by the helicopter, and the parallel operation rule conditions are satisfied, then obtain the corresponding installation loss according to the power of the power turbine and the power required by the helicopter under non-installed conditions.

2. The method for calculating the installation loss of a turboprop engine according to claim 1, characterized in that Before obtaining the second input data from a performance calculation model of a turboshaft engine established based on bench data, it includes: Selecting the numerical values of the dependent variables; Establishing a performance calculation model of a turboshaft engine based on bench data according to the numerical values of the dependent variables and the performance calculation method of the turboshaft engine.

3. The method for calculating the installation loss of a turboshaft engine according to claim 1, characterized in that Obtaining the second input data from a performance calculation model of a turboshaft engine established based on bench data, including: Obtaining the calculation input conditions from a performance calculation model of a turboshaft engine established based on bench data; Obtaining the second input data according to the calculation input conditions.

4. The method for calculating the installation loss of a turboshaft engine according to claim 1, characterized in that, Obtaining the first input data, including: Obtaining the outlet boundary conditions and the inlet boundary conditions; Calculating the first input data according to the outlet boundary conditions and the inlet boundary conditions.

5. The method for calculating the installation loss of a turboshaft engine according to claim 4, characterized in that, Obtaining the outlet boundary conditions and the inlet boundary conditions, including: Defining the inlet of the intake duct / compressor as the section of the turboshaft engine, and the corresponding first data; Defining the outlet of the gas turbine / inlet of the power turbine as the section of the turboshaft engine, and the corresponding second data; Setting the outlet boundary conditions and the inlet boundary conditions based on the first data and the second data.

6. The method for calculating the installation loss of a turboshaft engine according to claim 1, wherein The method further includes: If the conditions are not met: The target total temperature at the outlet of the gas turbine is equal to the second total temperature, the target total pressure at the outlet of the gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power of the power turbine is equal to the power required by the helicopter, re-execute step two until all conditions are met.

7. The method for calculating the installation loss of a turboprop engine according to claim 6, wherein Not meeting the conditions includes: Not meeting: the target total temperature at the outlet of the gas turbine is equal to the second total temperature, the target total pressure at the outlet of the gas turbine is equal to the second total pressure, then adjust the total temperature and the total pressure in the first input data until the conditions are met; Not meeting: The second flow rate is equal to the first flow rate, then adjust the static pressure in the first input data until the conditions are met; Not meeting: The power of the power turbine is equal to the power required by the helicopter, then adjust the compressor speed until the conditions are met.

8. The method for calculating the installation loss of a turboshaft engine according to claim 1, characterized in that, If the parallel operation rule conditions are met, it includes: The rotational speeds of the left and right corresponding turboshaft engine gas generators are equal, or the torques of the left and right corresponding turboshaft engine power turbines are equal. Otherwise, execute Step 2 until the conditions are met.

9. The method for calculating the installation loss of a turboshaft engine according to claim 1, characterized in that, If the conditions are met: the total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, and the power of the power turbine is equal to the power required by the helicopter, then the corresponding installation loss is obtained based on the power of the power turbine under non-installed conditions and the power required by the helicopter, including: Calculate the installation loss at constant rotational speed and the installation loss at constant temperature.

10. A device for calculating the installation loss of a turboshaft engine, characterized in that, Including: An acquisition module for obtaining second input data from a performance calculation model of a turboshaft engine established based on bench data, where the second input data includes the first flow rate at the inlet of the turboshaft engine section and the first total temperature and first total pressure at the inlet of the gas turbine; A first calculation module for obtaining first input data, where the first input data includes the second total temperature, second total pressure and second flow rate of the turboshaft engine section, as well as the power of the power turbine and the power required by the helicopter; A second calculation module for obtaining first sub-input data based on the first total temperature and first total pressure, where the first sub-input data includes the total temperature at the outlet of the target gas turbine and the total pressure at the outlet of the target gas turbine, and the first sub-input data is obtained by the same calculation method as the first input data; A condition module for, if the conditions are met: the total temperature at the outlet of the target gas turbine is equal to the second total temperature, the total pressure at the outlet of the target gas turbine is equal to the second total pressure, the second flow rate is equal to the first flow rate, the power of the power turbine is equal to the power required by the helicopter, and the conditions for the parallel operation rule are met, then obtaining the corresponding installation loss based on the power of the power turbine under non-installed conditions and the power required by the helicopter.

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