KR101409361B1 - Method for modeling engine performance for aircraft - Google Patents

Abstract

The present invention relates to a low-pass turbofan engine performance modeling technique for supersonic aircraft.
As a form for realizing this, the present invention provides a method of modeling an engine performance of an aircraft, comprising the steps of: (a) selecting a target engine to be modeled; (b) collecting basic performance data including values of design parameters of the target engine; (c) analyzing sensitivity to the design variables collected through the step (b) and extracting one or more important variables that affect the performance of the target engine; (d) determining an intake port voltage recovery rate of the target engine to be modeled; (e) determining a turbine cooling air flow rate of the target engine to be modeled; (f) determining a fan pressure ratio of the target engine to be modeled; And (g) setting the turbine cooling air flow rate determined from the step (e) and the fan pressure ratio determined in the step (f) as conditions for setting the intake port voltage recovery rate of the target engine determined from the step (d) And performing an optimization analysis to obtain an optimal objective function value from the one or more important design variables extracted in the step (c), wherein the optimization analysis in the step (g) is performed by an adaptive random search method Random Search Method). ≪ / RTI >

Description

METHOD FOR MODELING ENGINE PERFORMANCE FOR AIRCRAFT FIELD OF THE INVENTION [0001]

The present invention relates to a method of modeling an engine performance of an aircraft, and more particularly to a low-pass turbofan engine performance modeling technique for a supersonic aircraft.

Generally, gas turbine engine performance simulations are generated by gas turbine engine developers based on vast amounts of engine test data collected during engine development.

Detailed information about these simulations is kept secret by the engine developer, and only a few parameter values are accessible to buyers for control purposes.

Many researchers want information on the engine's performance and lifetime based on the gas turbine engine performance model, so they want information on the specific parameters of the engine they operate.

In the collaborative relationship of the engine development program, engine component designers and engine performance researchers are forced to develop gas turbine simulations based on limited information, since latecomers only have to acquire partial cycle data published by leading companies.

SUMMARY OF THE INVENTION The present invention has been made in view of the foregoing, and it is an object of the present invention to provide a low-pass turbofan engine for a supersonic aircraft equipped with a late burner, ). The purpose of the model is to model the performance of the system.

According to another aspect of the present invention, there is provided a method of modeling an engine performance of an aircraft, the method comprising: (a) selecting a target engine to be modeled; (b) collecting basic performance data including values of design parameters of the target engine; (c) analyzing sensitivity to the design parameters collected through the step (b), and extracting one or more important design variables that affect the performance of the target engine; (d) determining an intake port voltage recovery rate of the target engine to be modeled; (e) determining a turbine cooling air flow rate of the target engine to be modeled; (f) determining a fan pressure ratio of the target engine to be modeled; And (g) setting the turbine cooling air flow rate determined from the step (e) and the fan pressure ratio determined in the step (f) as conditions for setting the intake port voltage recovery rate of the target engine determined from the step (d) Performing an optimization analysis for each of the one or more critical design variables extracted in step (c) to find values in the design area that satisfy the engine performance of the aircraft for each of the one or more critical design variables Wherein the optimization analysis in the step (g) is performed by an adaptive random search method.

According to a preferred embodiment, the engine may be for a low-bypass turbofan engine for a supersonic aircraft.

According to a preferred embodiment, the intake port voltage recovery rate of the target engine may be a value obtained by applying a coefficient of 0.88 to 1.37 to the intake port voltage recovery rate diagram provided through the GASTURB program.

According to a preferred embodiment, the turbine cooling air flow rate of the target engine may be in the range of 15% to 25% of the engine core mass flow.

1 is a flow chart illustrating a method for modeling an engine performance of an aircraft according to the present invention;
2 is a view showing an F100-PW-229 low-bypass turbo fan engine, which is an example of a target engine applied to the modeling method according to the present invention.
3 is a view showing an engine model for cycle analysis of the F100-PW-229 low-bypass turbo fan engine of FIG. 2;
FIG. 4 is a view showing a sensitivity analysis result according to a change of 5% of main design parameters of the F100-PW-229 low-bypass turbofan engine of FIG. 2; FIG.
5 is a view showing the inlet voltage force recovery rate of the F100-PW-229 low-bypass turbo fan engine of FIG. 2;
6 is a view showing the number of compressors and the pressure ratio tendency at each stage of the F100-PW-229 low-bypass turbo fan engine of FIG. 2;
FIG. 7 is a diagram for explaining an adaptive random search method applied to optimization analysis of a modeling method according to the present invention. FIG.
Figs. 8 and 9 are diagrams showing the performance diagrams of the low-pressure compressor and the high-pressure compressor applied to the off-design point analysis, respectively.
10 is a view showing an engine model and a comparison value of a low-pressure compressor operating line of Reference 1;
11 is a diagram showing performance curves and operating lines of a high-pressure turbine applied to the modeling method according to the present invention.
12 is a diagram showing performance curves and operating lines of a low-pressure turbine applied to the modeling method according to the present invention.
FIGS. 13 to 16 are diagrams showing partial partial performance analysis results applied to the modeling method according to the present invention. FIG.
17 and 18 are graphs showing the total thrust force and the fuel flow rate according to the altitude and the speed change under the military power condition, respectively.
19 and 20 are graphs showing total thrust force and fuel flow rate according to altitude and speed change under the condition of maximum AB power, respectively.

Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings. In the drawings, the same reference numerals are used to designate the same or similar components throughout the drawings. In the following description of the present invention, a detailed description of known functions and configurations incorporated herein will be omitted when it may make the subject matter of the present invention rather unclear.

Meanwhile, as will be apparent from the following description, it should be understood that the present invention is also applicable to other aircraft engines, although the present invention is described by taking a low-bypass turbofan engine for a supersonic aircraft equipped with a later combustor as an example.

1 is a flowchart illustrating a method of modeling an engine performance of an aircraft according to the present invention.

Referring to FIG. 1, a method for modeling an engine performance of an aircraft according to the present invention includes: a first step (S101) of selecting a target engine to be modeled; A second step (S102) of collecting basic performance data including values of design parameters of the target engine; A third step (S103) of analyzing sensitivity to the design variables collected through the second step and extracting one or more important variables that affect the performance of the target engine; A fourth step (S104) of determining the intake port voltage recovery rate of the target engine to be modeled; A fifth step (S105) of determining a turbine cooling air flow rate of the target engine to be modeled; A sixth step (S106) of determining a fan pressure ratio of the target engine to be modeled; And the turbine cooling air flow rate determined from the fifth step, and the fan pressure ratio determined in the sixth step are set as conditions for setting the intake-port voltage recovery rate of the target engine determined from the fourth step, the turbine cooling- And a seventh step (S107) of performing an optimization analysis to obtain an optimal objective function value from the at least one important design variable.

According to a preferred embodiment, the optimization analysis in the seventh step (S107) may be performed by an adaptive random search method.

Referring to FIG. 1, details of each detailed step of a method for modeling an engine performance of an aircraft according to the present invention will be described.

[ Modeling  A first step of selecting a target engine to be a target S101 )]

The present inventors performed a process of selecting a target engine to be modeled as a first step of a method for modeling an engine performance of an aircraft according to the present invention.

According to a preferred embodiment, the target engine may be of a low bypass F100-PW-229 engine fitted with a late combustor and used in F-15 and F-16.

2 shows a F100-PW-229 low-bypass turbo-fan engine, which is an example of a target engine applied to the modeling method according to the present invention.

The F100-PW-229 engine has a two-spool turbo fan engine configuration consisting of a three-stage fan, a ten-stage compressor, a two-stage low pressure turbine and a two-stage high pressure turbine. In addition, the F100-PW-229 low-bypass turbo fan engine generates Military Power, 79.1kN, Maximum Afterburner Power, and 129.6kN of thrust. The non-fuel consumption rate is 77.5kg / kN · h, Maximum Afterburner Power, and 197.8 kg / kN · h.

[The second step of collecting the basic performance data including the values of the design parameters of the target engine S102 )]

As a second step of a method for modeling an engine performance of an aircraft according to the present invention, the inventors performed a process of collecting basic performance data including values of design parameters of a target engine through a website and a public route.

For example, the amount of data available to generate engine cycle simulations and their suitability vary from condition to condition.

The market brochures of the engines available to the public show only some key values, such as the engine's bypass ratio, turbine temperature, and flow rate, instead of the actual engine data. In addition, since the values of the other operating points are mixed and shown, reliability as the source data is low.

Table 1 below is the F100-PW-229 engine base data collected via the website and public route.

[Table 1]

In the method of modeling the engine performance of the aircraft according to the present invention, the engine performance modeling was performed based on the website and related related data. Table 1 shows basic performance data of the collected F100-PW-229 turbo fan engine Show. The contents of Table 1 do not show the specific value for the engine, and it is understood that the engine performance value should be predicted within a certain range.

Meanwhile, the following Table 2 shows the component characteristics of the F100-PW-229 engine collected through the website and public route. The characteristic value of the components of the engine is indispensably required in the analysis of the undesigned point, and the performance of the engine can be confirmed accurately according to the reliability thereof.

[Table 2]

A third step of analyzing the sensitivity of the design parameters collected through the second step and extracting one or more important variables that affect the performance of the target engine S103 )]

As a third step of the method for modeling the engine performance of an aircraft according to the present invention, the present inventors extracted one or more important variables to be applied to a simulator program (for example, a GASTURB program) used for low-bypass turbofan engine modeling .

3 shows an engine model for the F100-PW-229 low-bypass turbo fan engine cycle analysis. The inventors of the present invention found that the design parameters other than the collected design variables for the F100-PW-229 turbofan engine design point analysis Sensitivity analysis for thrust and fuel consumption rate was performed for design variables.

Figure 4 shows the sensitivity analysis results for 5% variation of major design variables.

4, Turbine Cooling Air, fan and turbine efficiency among the design variables have a great influence on engine performance.

[ Modeling  The inlet voltage of the target engine The recovery rate  The fourth step of determining S104 )]

As a fourth step of a method for modeling an engine performance of an aircraft according to the present invention, the present inventors determine the intake port voltage recovery rate of a target engine, which is a set value to be applied to a simulator program (for example, a GASTURB program) .

Specifically, the inlet voltage force recovery rate information for specific engines is extremely rare, and the present inventors have found that the inlet voltage regeneration rate basically provided in a simulator program (for example, a GASTURB program) used for engine modeling has a coefficient of 0.88 to 1.37 And the performance map was reconfigured so as to converge to the target values of the design point and the off-design point.

5 shows the intake port voltage recovery rate reconstructed for engine performance modeling according to the fourth step of the present invention.

[ Modeling  A fifth step of determining the turbine cooling air flow rate of the target engine to be the target ( S105 )]

As a fifth step of a method for modeling an engine performance of an aircraft according to the present invention, the inventors of the present invention calculate a turbine cooling air flow rate of a target engine, which is a set value to be applied to a simulator program (for example, a GASTURB program) .

For example, in order to set the turbine cooling air flow rate of the F100-PW-229 low-bypass turbo fan engine, it is necessary to set the life time of the turbine and the engine performance value in consideration of the value of the engine performance. The turbine cooling air flow rate was set within the range of 15% to 25% of the mass flow.

[ Modeling  Fans of the target engine to be targeted Pressure ratio  (Step < RTI ID = 0.0 > S106 )]

As a sixth step of a method for modeling an engine performance of an aircraft according to the present invention, the inventors of the present invention determine the fan pressure ratio of a target engine which is a set value to be applied to a simulator program (for example, a GASTURB program) .

For example, in the GASTURB program, the fan of the F100-PW-229 Turbo Fan Engine refers to the low pressure compressor and has an Inner fan area (Station 25) and an Outer fan area (station 13) .

For example, FIG. 6 shows the number of compressors and the pressure ratio tendency at each end, where each point represents the pressure ratio of the fan and high pressure compressor applied to the analysis. The core and bypass pre-pressure ratio should closely match to obtain maximum thrust. The third-stage fan compression ratio was set to 4.04, and the voltage outputs of the core and the bypass outlet were 327.71 kPa and 287.71 kPa, respectively. The fan pressure ratio and the pressure ratio of the high- Able to know.

[From the one or more critical design variables extracted in the third step, Function value A seventh step of performing optimization analysis for obtaining S107 )]

The seventh step of the method for modeling an engine performance of an aircraft according to the present invention is a method for calculating an optimized objective function value based on the data calculated in the fourth to sixth steps The process of minimizing "Total Fuel Flow" was performed.

According to a preferred embodiment, the optimization analysis process as a seventh step of the present invention includes the steps of: calculating a recovery rate of the intake port voltage of the target engine determined from the fourth step, the turbine cooling air flow rate determined from the fifth step, The fan pressure ratio may be set as a set condition.

ㆍ Optimization theory

The low-bypass turbofan engine optimization problem is a global optimization problem in which there are several local extreme points, so a large amount of computation is required to find the optimum point.

For example, the Random Search Method is one of the widely used methods for global optimization problems. This method has many advantages when the algorithm is relatively simple and has little effect on the shape of the objective function, so that information on the characteristics of the objective function is applied to the engineering problem which is extremely limited.

However, the convergence of the random search method can only be referred to in terms of probability, and in a strict sense convergence can be ensured only when the number of calculations of the objective function becomes infinite.

Also, convergence of this random search method is directly affected by the mean value and variance of the random trial number directly. Therefore, in the present invention, an adaptive random search method is used in which a part of the random search method is modified to improve convergence speed.

The adaptive random search method is a method that corrects the average value and variance of random estimates for each optimization step based on the route shown in the optimization process so far.

Assuming that the number of variables to be optimized is "n", the random estimation variable is generated with a uniform probability in the n-dimensional sphere as shown in the following equation (1). This relationship is shown in Fig.

In the above equation (1), V _i ^* and R _i are n-dimensional vectors, which means the optimum design parameters up to the present and the tracking direction.

The variable "c" means the radius of the search sphere, and "k"

If R _i is 0, the adaptive random search method finds the optimal point of the next step only around the optimum point that has been traced so far. The following formula (2) is used to determine the tracking range of the next step from the minimization process so far.

The variables applied to the equations (1) and (2) are defined as follows.

The following Tables 3 and 4 show the main parameters and constraint conditions for the analysis satisfying the military power, respectively.

[Table 3]

[Table 4]

Table 5 below shows parameter optimal results for the F100-PW-229 low-bypass turbo-fan engine.

In Table 5, the target value of the design point is set to Military Power.

[Table 5]

Here, Military Power means the maximum usable thrust when the late combustor is not operating. In this case, Sea level, Mach number 0.2, ISA, Bleed Air 0%, Power offtake 12kW were analyzed.

The following Table 6 shows the results of the design point analysis.

[Table 6]

As can be seen from Table 6, it can be seen that the result value satisfying the requirement of the design point analysis result for the military power is calculated.

[Engine performance of aircraft Modeling  Other processes for

◎ Off-design point  Translate

In addition to the design point analysis by the process of the first step (S101) to the seventh step (S107) of the present invention as described above, Can be more accurately modeled.

Specifically, proper component design is required to predict aircraft component efficiency at points outside the design point. A general way to accumulate component performance leads is to multiply the performance data of the performance lead by the scale factor value to fit the design point data of the new engine and multiply the whole data of the component performance. This method is capable of creating a reliable component lead by matching the performance of the existing engine with that of the new engine. It is very difficult to obtain a performance map similar to the actual performance. The present inventors applied the accumulation technique applied in the GASTURB program as follows to accumulate the similar component performance diagram and apply it to the analysis.

The following Table 7 shows the scale factor of the low, high-pressure compressor, turbine performance diagram applied to the engine performance model calculated using the above equations (3) to (6).

[Table 7]

ㆍ Part load performance analysis

Although the value of the design point is accurate in the design point analysis, accurate engine performance may not be possible if the characteristic value of the component applied in the design point analysis does not reflect the characteristics of the actual engine. The characteristics of the engine components are unique to the engine manufacturer, so there is no data available. Therefore, the performance of compressors similar to those of the F100-PW-229 turbofan engine compressor was investigated. B.L. Koff, "F100-PW-229 Higher Thrust in Same Frame Size", J. Eng. A compressor capable of applying compressor performance data through the engine test data described in the literature of Gas Turbines Power, Volume 111, Issue 2, April, 1989 (hereinafter referred to as Reference 1) Performance graphs were generated. The values between the test data were constructed by interpolation. Figs. 8 and 9 show performance diagrams of a low-pressure compressor applied to off-design point analysis.

10 shows the engine model and the low pressure compressor operating line comparison value of the reference 1. It can be seen from FIG. 10 that the analysis results and the test data fit well.

Also, because the operating range of the turbine is very narrow for the turbine performance diagram, the general turbine performance diagram provided by the GASTURB program is scaled and applied to the analysis.

11 and 12 show a turbine performance diagram and an operating line applied to the engine model. Specifically, FIG. 11 shows the performance line and the operating line of the high-pressure turbine, and FIG. 12 shows the performance line and the operating line of the low-pressure turbine.

The partial load performance is defined as the performance according to the number of revolutions of the engine gas generator which is less than the design required correction rotational speed.

13 to 16 show partial load performance analysis results. Specifically, FIG. 13 shows the result of "Gross Thrust vs. GG Spool Speed", FIG. 14 shows the result of "Fuel Flow vs. GG Spool Speed", FIG. 15 shows the result of "Mass Flow vs. GG Spool Speed" , And Fig. 16 shows the result of "HPC pressure ratio versus GG spool speed ".

ㆍ Analysis according to altitude and speed change

FIGS. 17 and 18 show the gross thrust and the fuel flow according to the altitude, the speed change, and the curves show the engine deck data and the analysis result values from Sea Level to 10kft, 30kft, 50kft, and 60kft from the top to the bottom. Specifically, FIG. 17 shows the results of "Gross Thrust vs. Mach Number (Military Power)", and FIG. 18 shows the results of "Fuel Flow vs. Mach Number (Military Power)".

Referring to FIGS. 17 and 18, Gross Thrust shows a difference in sea level and supersonic region, but in other regions, the result of the engine model is similar to that of the engine deck. The fuel consumption rate increases as the speed increases, but the tendency is similar. The reason for this difference is the accuracy of the component values applied as the estimates. In particular, since the inlet pressure recovery rate directly affects the engine inflow flow rate, the values applied to the analysis are estimated from arbitrary values, This is a major cause. It is understood that more reliable analytical value can be obtained if the detailed performance limit of the actual engine and the component performance are applied to the engine performance model constructed through this process.

◎ Cycle analysis of late combustor application

High-performance military supersonic aircraft are not always flying at high speeds and should be capable of increasing thrust only for a short period of time if needed. Therefore, the latest fighter is equipped with a late burner in a low-bypass turbofan engine superior to a turbojet engine.

Based on the F100-PW-229 engine performance model based on the Military Power standard, the performance model was reconstructed with the engine performance model considering the late combustor.

ㆍ Application of late combustor Design point  Translate

In order to converge the performance value of the engine using the late burner through the engine deck, the design parameters of the late combustor were set by using the GASTURB program parametric iterative analysis method.

In Sea Level, ISA condition, the engine performance value of late combustor is 113.727kN of Gross Thrust and 6.7kg / s of fuel amount.

The following Table 8 shows the values of the late combustor design variables set through the parametric iterative analysis.

[Table 8]

Theoretically, it is known that the combustion temperature of the latter combustor reaches an approximate stoichiometry of 2300K to 2500K, assuming that all the oxygen is consumed.

The combustion temperature of the latter combustor can be predicted by using the following equation (7), and it is predicted that the combustion temperature of the latter combustor can be predicted using the following equation (7) It can be seen that it can reach about 2337K by applying the late combustor inlet temperature of 870K and stoichiometric temperature of 2500K given in Exhibit, Aug, 2-5, 2009 (hereinafter referred to as Reference 2).

In the engine performance model of the latter combustor, 2200 K was applied to the combustion temperature of the latter combustor in consideration of the loss.

Since the efficiency of the latter combusts the air mixed with the combusted gas, the combustion efficiency is lower than the combustor efficiency (0.98).

As a result of analysis of the design point for the late combustor, it can be understood that the result of analysis of the late combustor inlet temperature is 864K, and the analysis result of the engine applied to the late combustor is 1.16% for the gross thrust error and 0.93% for the fuel flow error. Able to know.

ㆍ Application of late combustor Off-design point  Translate

19 and 20 show the gross thrust and the fuel flow according to the altitude and the velocity change. 19 shows the result of "Gross Thrust vs. Mach Number (MAX. Afterburner Power)", and FIG. 20 shows the result of "Fuel Flow vs. Mach number (MAX.

In FIGS. 19 and 20, the curves show engine deck data and analysis results from Sea Level to 10kft, 30kft, 50kft, and 60kft from top to bottom.

Gross Thrust shows a similar tendency to the engine deck results, but the fuel consumption rate differs in the low and supersonic regions. The difference is that the fuel consumption rate in the latter combustor is proportional to the fuel consumption rate in the combustor plus the inflow air flow rate, so that the inlet pressure recovery rate applied to the analysis can be considered to be large.

The foregoing description is merely illustrative of the technical idea of the present invention, and various changes and modifications may be made by those skilled in the art without departing from the essential characteristics of the present invention. Therefore, the embodiments disclosed in the present invention are intended to illustrate rather than limit the scope of the present invention, and the scope of the technical idea of the present invention is not limited by these embodiments. The scope of protection of the present invention should be construed according to the following claims, and all technical ideas falling within the scope of the same shall be construed as falling within the scope of the present invention.

Claims (4)

A method of modeling an engine performance of an aircraft,
(a) selecting a target engine to be modeled;
(b) collecting basic performance data including values of design parameters of the target engine;
(c) analyzing sensitivity to the design parameters collected through the step (b), and extracting one or more important design variables that affect the performance of the target engine;
(d) determining an intake port voltage recovery rate of the target engine to be modeled;
(e) determining a turbine cooling air flow rate of the target engine to be modeled;
(f) determining a fan pressure ratio of the target engine to be modeled; And
(g) determining a turbine cooling air flow rate determined from the step (e), and the fan pressure ratio determined in the step (f) as conditions for setting the intake port voltage recovery rate of the target engine determined from the step (d) performing an optimization analysis for each of the one or more critical design variables extracted in step (c) to find values within the design area that satisfy the engine performance of the aircraft for each of the one or more critical design variables, ,
Wherein the optimization analysis in step (g) is performed by an adaptive random search method. The method according to claim 1,
Wherein the engine is a low-bypass turbo-fan engine for a supersonic aircraft. 3. The method of claim 2,
Wherein the intake port voltage recovery rate of the target engine is a value obtained by applying a coefficient of 0.88 to 1.37 to a suction port voltage recovery rate diagram provided through the GASTURB program. 3. The method of claim 2,
Wherein the turbine cooling air flow rate of the target engine is in the range of 15% to 25% of the engine core mass flow.

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