CN113378328B - Gas turbine front temperature calculation method for control system - Google Patents
Gas turbine front temperature calculation method for control system Download PDFInfo
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- CN113378328B CN113378328B CN202110756960.0A CN202110756960A CN113378328B CN 113378328 B CN113378328 B CN 113378328B CN 202110756960 A CN202110756960 A CN 202110756960A CN 113378328 B CN113378328 B CN 113378328B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
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Abstract
The invention discloses a method for calculating the front temperature of a gas turbine for a control system, which relates to the technical field of gas turbinesGeneral assemblyAnd power turbine rear gas specific enthalpy H5, calculating power turbine front gas specific enthalpy H45, calculating power turbine front temperature T45, calculating standard natural gas turbine front temperature T41 from T45Sign boardFrom T41Sign boardCalculating a gas turbine front temperature T41, determining a temperature correction value TtrimFrom TtrimCalculating a corrected gas turbine front temperature T41trimThe above functional relationship and the correlation coefficient are determined by the engine characteristics and written into the memory of the electronic controller in the form of a data interpolation table, and by the calculation method of the invention, the control system can obtain the accurate gas turbine front temperature.
Description
Technical Field
The invention relates to the technical field of gas turbines, in particular to a method for calculating the temperature of a gas turbine in front of a control system.
Background
The gas turbine of the aircraft engine is a key hot-end component of the engine subjected to high rotating speed and high temperature and is a key component for limiting the service life of the engine, and the temperature before the gas turbine is a key factor for influencing the service life of the gas turbine.
A typical free-turbine turboprop model is schematically illustrated in fig. 2. to monitor and control the gas turbine front temperature, the engine control system should obtain the T41 temperature (gas temperature in section 41), which is typically measured using a thermocouple. However, conventional thermocouple measurement devices are difficult to withstand due to the generally high T41 (greater than 1200 ℃). Therefore, the prior art often indirectly monitors the gas turbine front temperature by measuring and monitoring T45 (gas temperature in section 45) or T5 (gas temperature in section 5).
The prior art has the following disadvantages:
the power turbine front temperature T45 and the gas turbine front temperature T41 have a determined corresponding relation, T41 can be indirectly monitored to a certain extent by monitoring T45, but in a full-envelope range, the T45 cannot accurately reflect the temperature level of T41, and a thermocouple is arranged in a power turbine channel to bring aerodynamic loss, and if the thermocouple breaks, the damage to turbine parts is caused, so that the reliability and the safety of the whole machine are affected. The relationship between exhaust temperatures T5 and T41 is complex and simply monitoring T5 is not sufficient to ensure gas turbine safety.
Disclosure of Invention
The invention aims to provide a gas turbine front temperature calculation method for a free turbine type turboprop engine control system, which is applied to a digital electronic controller and used for monitoring the gas turbine front temperature of the free turbine type turboprop engine in real time so as to overcome the defects of the prior art.
The purpose of the invention can be realized by the following technical scheme:
a gas turbine front temperature calculation method for a control system, comprising the steps of:
s1: the electronic controller measures, collects and calculates required parameters, and the parameters comprise: the engine rotor speed Ng (the low-pressure rotor speed NL of the double-rotor engine), the power turbine speed NP, the engine output power HP, the air inlet temperature T2, the atmospheric static pressure P0, the dynamic pressure DP and the exhaust temperature T5 are sent to a control system;
s2: calculating power gas flow W by Ng, T2, standard natural gas compressor inlet flow function and gas flow functiong45From HP and T5, the total power HP of the power turbine is calculated respectivelyGeneral assemblyAnd power turbine after-gas specific enthalpy H5, according to Wg45、HPGeneral assemblyAnd H5, calculating to obtain the specific enthalpy H45 of the gas before the power turbine, wherein the functional relation and the related coefficient are determined by the characteristics of the engineSex determination, which is obtained by a design and test mode and written into a memory of the electronic controller in the form of a data interpolation table;
s3: calculating the power turbine front temperature T45 from H45;
s4: standard natural gas turbine front temperature T41 calculated from T45Sign boardFrom T41Sign boardCalculating a gas turbine front temperature T41;
s5: for each engine, the accurate gas turbine front temperature is obtained through the actual measurement or calculation of the vehicle platform, and is compared with the calculated T41 to determine the temperature correction value TtrimFrom TtrimCalculating a corrected gas turbine front temperature T41trim。
Further: in S2, the standard natural gas compressor inlet flow function is WSign boardF (NgC, H), the gas flow function is Wg45＝f(WInto) Wherein, in the step (A),
p2 is obtained by converting P0 and DP, and H is the flying height and is obtained by converting P0.
Further: the calculation formula of H45 is H45 ═ HPGeneral assembly/(Wg45X η) + H5, wherein HPGeneral assembly＝HP/ηMachine for working，H5＝T5×Cpg，ηMachine for workingFor mechanical efficiency, CpgIs the constant pressure specific heat capacity of the fuel gas, and eta is the turbine efficiency.
Further: in the S3-S7, T45 is H45/Cpg，T41Sign boardT45 × k, k is the temperature ratio coefficient of 41 cross section to 45 cross section, T41 is f (T41)Sign board)，T41trim＝T41+Ttrim。
Further: the functional relation and the correlation coefficient are determined by engine characteristics, are obtained by a design and test mode, and are written into a memory of the electronic controller in the form of a data interpolation table.
The gas turbine front temperature calculation related input parameters are given in the following table:
| name (R) | (symbol) | Source |
| Rotational speed of output shaft | NP | Rotating speed sensor |
| Output power of engine | HP | Torque sensor |
| Low rotor speed | NL | Rotating speed sensor |
| Temperature of inlet air | T2 | Temperature sensor |
| Flying speed | TAS | Aircraft with a flight control device |
| Static pressure of atmosphere | P0 | Pressure sensor |
| Dynamic pressure | DP | Aircraft with a flight control device |
| Exhaust temperature | T5 | Exhaust gas temperature sensor |
| Gas turbine front temperature correction | trim |
Parameters such as the rear exhaust temperature (T5) of the power turbine, the rotating speed Ng of the gas generator, the output power HP of the engine and the like are measured through an electronic controller, and the parameters are obtained through calculation and correction, so that the temperature level of the gas turbine component is reflected.
The method is simple and practical, has wide application range, can realize the accurate monitoring of the control system on the temperature in front of the gas turbine, is applied to an electronic controller of an aviation turboprop engine at present, and has reasonable and feasible test results.
The invention has the beneficial effects that:
the parameters required by the measurement and acquisition calculation of the electronic controller comprise: the method comprises the steps that the engine rotor rotating speed Ng (the low-pressure rotor rotating speed NL of a double-rotor engine), the power turbine rotating speed NP, the engine output power HP, the air inlet temperature T2, the atmospheric static pressure P0, the dynamic pressure DP and the exhaust temperature T5 are sent to a control system, and the total power HP of the power turbine is respectively calculated according to the inlet flow function and the gas flow function of a standard natural gas compressorGeneral assemblyAnd the power turbine rear gas specific enthalpy H5, calculating to obtain the power turbine front gas specific enthalpy H45, calculating the power turbine front temperature T45, and calculating the standard natural gas turbine front temperature T41 from T45Sign boardFrom T41Sign boardCalculating a gas turbine front temperature T41, determining a temperature correction value TtrimFrom TtrimCalculating a corrected gas turbine front temperature T41trimBy the calculation method of the invention, the control system can obtain accurate gas turbine front temperatureAnd (4) degree.
Drawings
In order to facilitate understanding for those skilled in the art, the present invention will be further described with reference to the accompanying drawings.
FIG. 1 is a schematic diagram of the calculation process of the present invention;
FIG. 2 is a component level model schematic of a typical turboprop engine.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to fig. 1-2, a preferred embodiment of the present invention provides a gas turbine front temperature calculation method for a control system, S1: the electronic controller measures, collects and calculates required parameters, and the parameters comprise: the engine rotor speed Ng (the low-pressure rotor speed NL for the twin-rotor engine), the power turbine speed NP, the engine output HP, the intake air temperature T2, the atmospheric static pressure P0, the dynamic pressure DP, and the exhaust gas temperature T5 are sent to the control system.
S2: A. environment and conversion parameter analysis:
reduced speed of low pressure rotor
Total pressure of intake air PGeneral assemblyThe intake passage loss coefficient x (P0+ DP) and the intake total pressure coefficient θ PGeneral assembly/101.325kpa；
Temperature coefficient of intake air
The temperature difference between the local environment static temperature and the standard atmospheric environment static temperature is as follows:
△T0＝T2–TAS2/(2Cp)-(288.135–6.5×H)
wherein TAS (m/s) is vacuum speed, Cp is air constant pressure specific heat capacity, and H is altitude, and the unit is km.
B. Calculating the gas flow of the power turbine:
power turbine gas flow Wg45The air-conditioning system is obtained by calculation of parameters such as NLC, height H, air-entraining mode, air-intake pressure, air-intake temperature T2 and the like, and specifically comprises the following steps:
1. calculating the reference flow W of the working line of the low-pressure compressor by means of NLC and height H interpolationDatum(kg/s) obtained by interpolation of compressor characteristics;
2. reference flow W on working line of low-pressure compressorDatumOn the basis of (kg/s), the intake temperature and the intake pressure are converted and relevant correction is carried out to obtain the actual total flow W of the low-pressure engineGeneral assembly：
WGeneral assembly＝WDatum×γ×θ×k1×k2
Wherein gamma is an intake air temperature coefficient,
theta is an intake pressure coefficient, and theta is PGeneral assembly101.325 kpa; k1 is a vacuum speed and altitude correction coefficient obtained by interpolation of the vacuum speed and altitude; k2 is a Bleed air correction coefficient composed of NLC and Bleed air Ratio Bleed Ratio (Ratio of Bleed air amount to actual total compressor flow rate Bleed Ratio/W)General assembly) Obtaining by interpolation;
3. actual total flow W of low-pressure air machineGeneral assemblyOn the basis of the method, the power turbine gas flow W is obtained by considering the bleed air factor lead-in coefficient k3g45：
Wg45＝WGeneral assemblyXk 3, k3 from the actual total flow WGeneral assemblyAnd the Bleed air Ratio Bleed Ratio is obtained by interpolation.
C. Calculating the total work of the power turbine:
total power HP of power turbineGeneral assemblyThe power gain to compensate for the loss on the basis of the engine output power HP is calculated as follows:
HPgeneral assembly＝HP×k4
In the formula, k4 is the power turbine efficiency, which is obtained by interpolation of the power turbine characteristics.
D. Calculating specific enthalpy of gas after power turbine:
the specific enthalpy of combustion H46(kJ/kg) after the power turbine is directly obtained by interpolation of the exhaust temperature T6 and the difference delta T0 between the static temperature of the local environment and the static temperature of the standard atmospheric environment.
E. Calculating specific enthalpy of gas before power turbine:
the following equation can be obtained from the conservation of energy at the inlet and outlet of the power turbine:
Wg45×η×(H45–H46)＝HPgeneral assembly
Available power turbine front gas specific enthalpy H45(kJ/kg) calculation formula:
H45＝HPgeneral assembly/(Wg45×η)+H46
In the formula, Wg45For power turbine inlet gas flow, HPGeneral assemblyH46 is the specific enthalpy of the gas after the power turbine, and eta is the turbine efficiency.
S3: the total temperature T45(K) before the power turbine is directly obtained by interpolation of the specific enthalpy H45 of gas before the power turbine and the temperature difference DeltaT 0 between the static temperature of the local environment and the static temperature of the standard atmospheric environment.
S4: calculating the front total temperature of the high-pressure turbine:
1. the standard high-pressure turbine front total temperature T41 is marked as the product of T48 and the temperature ratio coefficient of the inlet section of the high-pressure turbine bucket and the inlet section of the power turbine, namely:
t41 scale T48 xk 5
In the formula, T48 is the total temperature before the power turbine; k5 is a temperature ratio coefficient of the inlet section of the high-pressure turbine bucket to the inlet section of the power turbine, and is obtained by interpolation of T48 and height H;
2. the method utilizes an NP correction coefficient k5, a temperature difference delta T0 correction coefficient k6, a bleed air correction coefficient k7 and an airspeed correction coefficient k8 to correct the front total temperature T41 of the standard high-pressure turbineSign boardAnd correcting to obtain an actual front total temperature T41 of the high-pressure turbine:
T41＝T41sign board×k6×k7×k8
In the formula, K6 is obtained by NP interpolation, K7 is obtained by T45 modification and temperature difference delta T0 interpolation, and K8 is obtained by Bleed Ratio interpolation of the Ratio of the air-entraining amount to the actual total flow of the compressor.
S5: gas turbine front temperature calculation:
the gas turbine front temperature is obtained by T41 calculation, and the calculation formula is as follows:
gas turbine front temperature T41trimThe temperature before the gas turbine is T41(K) + trim-273, wherein the temperature before the gas turbine trim is a corrected value of the temperature before the gas turbine, and the corrected value of each engine is determined through a correction program.
The foregoing is merely exemplary and illustrative of the present invention and various modifications, additions and substitutions may be made by those skilled in the art to the specific embodiments described without departing from the scope of the invention as defined in the following claims.
Claims (5)
1. A method of gas turbine pre-temperature calculation for a control system, comprising the steps of:
s1: the electronic controller measures, collects and calculates required parameters, and the parameters comprise: the engine rotor rotating speed Ng, the low-pressure rotor rotating speed NL of the double-rotor engine, the power turbine rotating speed NP, the engine output power HP, the air inlet temperature T2, the atmospheric static pressure P0, the dynamic pressure DP and the exhaust temperature T5 are sent to a control system;
s2: calculating to obtain gas flow W from Ng, T2, standard natural gas compressor inlet flow function and gas flow functiong45From HP and T5, the total power HP of the power turbine is calculated respectivelyGeneral assemblyAnd power turbine after-gas specific enthalpy H5, according to Wg45、HPGeneral assemblyH5, calculating to obtain the front gas specific enthalpy H45 of the power turbine;
s3: calculating the power turbine front temperature T45 from H45;
s4: standard natural gas turbine front temperature T41 calculated from T45Sign boardFrom T41Sign boardCalculating a gas turbine front temperature T41;
s5: for each engine, the accurate gas turbine front temperature is obtained by the actual measurement or calculation of the vehicle platform and is compared with the calculated T41 to determine the temperature correction valueTtrimFrom TtrimCalculating a corrected gas turbine front temperature T41trim。
2. The method of calculating the pre-turbine temperature of the gas turbine for the control system according to claim 1, wherein in the step S2, the standard natural compressor inlet flow function is WSign boardF (NgC, H), the gas flow function is Wg45＝f(WInto) Wherein, in the step (A),
NgC is the equivalent of Ng in the standard state, P2 is obtained by converting P0 and DP, and H is the flying height and is obtained by converting P0.
3. The method of claim 1, wherein the H45 calculation formula is H45 ═ HPGeneral assembly/(Wg45×η)+H5，HPGeneral assembly＝HP/ηMachine for working，H5＝T5×Cpg，ηMachine for workingFor mechanical efficiency, CpgIs the constant pressure specific heat capacity of the fuel gas, and eta is the turbine efficiency.
4. The method of claim 1, wherein T45 is H45/Cp in S3-S5g，CpgT41 being the constant pressure specific heat capacity of the gasSign boardT45 × k, k is the temperature ratio coefficient of 41 cross section to 45 cross section, T41 is f (T41)Sign board)，T41trim＝T41+Ttrim。
5. The method of calculating the pre-gas turbine temperature for a control system according to any one of claims 1 to 4, wherein the above functional relationships and correlation coefficients are determined by engine characteristics, obtained by design and experimental means, and written in the form of data interpolation tables into the memory of the electronic controller.
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| EP2489859A1 (en) * | 2011-02-18 | 2012-08-22 | Pratt & Whitney Canada Corp. | Free gas turbine with constant temperature-corrected gas generator speed |
| CN108843451A (en) * | 2018-05-31 | 2018-11-20 | 中国航发沈阳发动机研究所 | Gas-turbine combustion chamber outlet temperature calculation method |
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