Aerospace Engine High-Temperature Combustion Gas Temperature Inversion Method

Abstract

A method for inverting the high-temperature gas temperature of an aero-engine, characterized by comprising the following steps: S1, obtaining the radiation spectrum of the high-temperature gas through an aero-engine high-temperature gas spectral acquisition device; S2, determining the characteristic absorption peak band by analyzing the spectral characteristics of the high-temperature gas; S3, establishing a dual-band gas radiation transfer model according to the characteristic absorption peak band of the high-temperature gas; S4, constructing a high-temperature gas temperature solution deviation function and determining the constraint conditions; S5, solving the deviation function within the constraint conditions by the Levenberg-Marquardt algorithm to obtain N deviation function values. S6, taking the temperature T corresponding to the minimum deviation function value among the N deviation function values gas as the actual temperature value of the high-temperature gas. Compared with the traditional gas temperature measurement method, the present invention has the advantages of good measurement real-time performance, simple modification of the existing engine, simple measurement device, etc., reduces the application cost, and improves the stability of the measurement system.

Description

Technical Field

[0001] The invention relates to the field of optical technology, and in particular to a method for inverting the temperature of high-temperature combustion gas of an aero-engine. Background Art

[0002] With the development of my country's aviation industry, the demand for a new generation of advanced aircraft engines has further increased. With the development of the 14th Five-Year Plan and the formulation of the Outline of the 2035 Vision Goals, there is an urgent need for military and civilian high-performance aircraft engines. The development of a new generation of aircraft engines is imminent and cannot be delayed. In the development process of a new generation of aircraft engines, higher requirements are placed on the engine thrust-to-weight ratio and combustion efficiency, which has promoted the demand for monitoring the distribution of high-temperature gas combustion products in aircraft engines. The generation of high-temperature gas is closely related to the combustion of aviation kerosene during the operation of aircraft engines, and the distribution of high-temperature combustion products affects the combustion efficiency of aircraft engines. Therefore, temperature monitoring of high-temperature gas in aircraft engines is of great significance for evaluating the health of the engine and improving the operating life, reliability and power level of aircraft engines.

[0003] At present, there are mainly the following methods for measuring the temperature of aircraft engine gas:

[0004] (1) Direct thermocouple temperature measurement method: Place the thermocouple directly in the gas flow, so that the gas flow is in direct contact with the thermocouple. When the thermocouple reaches thermal equilibrium, the temperature indicated by the thermocouple is the temperature of the gas flow at that time.

[0005] (2) Probe sampling method: A method of analyzing the content of various components in the fuel gas and estimating the fuel gas temperature based on the fuel gas combustion chemical equation. The probe is used to extract a mixture of fuel gas and fuel, which is then quenched and transmitted to a gas analyzer through a pipeline to analyze the composition and content of the sample. The combustion efficiency and residual gas coefficient are calculated based on the combustion equation to estimate the fuel gas temperature. This method is often used to measure the distribution of the combustion chamber outlet temperature field during ground testing of aircraft engines.

[0006] (3) Tunable laser absorption spectroscopy technology: The gas temperature is measured by using the characteristic that the gas absorption spectrum changes with temperature. The line intensity of the specific gas absorption spectrum line has a functional relationship with the gas temperature. The absorption lines of the tunable laser absorption spectrum in two different wavelength absorption regions under the same absorption path are measured at the same time. Based on Beer-Lambert's law, the ratio function of the line intensity of the spectrum line corresponding to the two wavelengths and the gas temperature can be calculated to obtain the measured gas temperature;

[0007] (4) Fourier transform spectroscopy: Based on the fact that the rotational constants of gas molecules are different in different band emission bands, the temperature of the measured gas is calculated by establishing the relationship between the rotational constant of the molecule and the temperature. Specifically, a Fourier transform spectrometer is used to measure the spectral absorption lines of the hot gas. By measuring the gas radiation spectra at two different reference temperatures of the gas, the functional relationship between the rotational constant and the temperature in its rotational-vibrational spectral band can be obtained, and thus the temperature of the measured gas can be obtained.

[0008] In terms of the current development status at home and abroad, although the related technologies and equipment for infrared radiation temperature measurement have made great progress, most measurement systems still have technical problems such as the complex shape, high-speed rotation, and difficult positioning of turbine blades, which cannot be directly applied to the temperature measurement of turbine blades of aero-engines. There is no mature product in China yet. The existing technologies have the following disadvantages:

[0009] (1) Gas analysis measurement method: Since the gas sampling method is used to measure the gas temperature, the sampling probe has a complex structure. In order to prevent the sampling probe from being ablated by high-temperature gas, the sampling probe requires a large water-cooling mechanism for cooling, and the volume structure is relatively large. It can only be used for the ground experimental platform of aero-engines and cannot realize the on-line measurement of real-time gas temperature;

[0010] (2) Tunable diode laser absorption spectroscopy: Since the active laser measurement method requires two openings on the engine casing wall at the emission end and the receiving end of the laser, the modification of the existing engine is difficult. The application of the active technology is limited in the very small internal space of the engine. At the same time, due to the laser emission end and the receiving end being easily contaminated by the gas flow and carbon black particles during operation, the measurement accuracy is reduced;

[0011] (3) Fourier transform spectroscopy: Due to the unique structure of the Fourier transform spectrometer, the spectrum needs to be obtained by controlling the movement of the moving mirror. Therefore, it takes about 10 s to obtain the spectrum, which has the disadvantage of lag in the temperature measurement of high-speed gas flow and cannot reflect the real-time temperature of the engine. At the same time, the Fourier transform spectrometer is usually used to measure the tail flame emitted by the engine in a remote sensing measurement method and cannot measure the temperature of the gas inside the engine. It is very difficult to integrate the Fourier transform spectrometer with the aero-engine. Summary of the Invention

[0012] In view of the above problems, the object of the present invention is to propose a method for inverting the high-temperature gas temperature of an aero-engine, which realizes the real-time measurement of the gas radiation spectrum inside the aero-engine through a gas spectrum acquisition device, and proposes a passive spectral dual-wavelength constraint method based on the constraint condition that the measured radiation spectrum and the gas temperature remain unchanged during the measurement process.

[0013] This method does not require an active light source. By measuring the internal radiation spectrum distribution during the operation of an aeroengine, the gas temperature is inversely calculated. Compared with traditional gas temperature measurement methods, it has advantages such as good measurement real-time performance, simple modification of existing engines, and simple measurement devices, reducing the application cost and improving the stability of the measurement system.

[0014] To achieve the above object, the present invention adopts the following specific technical solutions:

[0015] The present invention provides a method for inversely calculating the high-temperature gas temperature of an aeroengine, which is realized based on a high-temperature gas concentration and temperature measurement system for an aeroengine, and includes the following steps:

[0016] S1. Obtain the radiation spectrum of the high-temperature gas through a high-temperature gas spectrum acquisition device for an aeroengine;

[0017] S2. Determine the characteristic absorption peak band by analyzing the spectral characteristics of the high-temperature gas;

[0018] S3. Establish a dual-band gas radiation transfer model according to the characteristic absorption peak band of the high-temperature gas:

[0019]

[0020] Among them,

[0021] λ1 and λ2 are respectively two narrow bands within the characteristic absorption band of the high-temperature gas;

[0022] τ gas (λ i ,T gas ) is the transmittance of the high-temperature gas;

[0023] T gas and T b are respectively the temperature of the high-temperature gas and the temperature of the turbine blade;

[0024] E total (λ1) is the total radiation energy received by the spectral measurement device at the characteristic band λ1;

[0025] E total (λ2) is the total radiation energy received by the spectral measurement device at the characteristic band λ2;

[0026] Δλ1 is the integration bandwidth at the characteristic band λ1;

[0027] Δλ2 is the integration bandwidth at the characteristic band λ2;

[0028] c1 = 3.7418×10 -16 W·m 2 is the first radiation constant;

[0029] c2 = 1.4388×10 -2 m·K is the second radiation constant;

[0030] S4. Construct a deviation function for solving the high-temperature gas temperature and determine the constraint conditions;

[0031] The deviation function is:

[0032]

[0033] The constraint conditions of the deviation function are:

[0034]

[0035] S5. Solve the deviation function within the constraint conditions through the Levenberg-Marquardt algorithm to obtain N deviation function values.

[0036] S6. Take the temperature T corresponding to the minimum deviation function value among the N deviation function values gas as the actual temperature value of the high-temperature gas.

[0037] Preferably, in step S3:

[0038] At the characteristic absorption band of the high-temperature gas spectrum, the radiation energy received by the high-temperature gas spectrum acquisition device of the aero-engine is:

[0039] L total = τ gas (λ i , T gas ) L b (λ i , T b ) + [1 - τ gas (λ i , T gas )] L(λ i , T gas ) (1 - 1)

[0040] Where

[0041] τ gas (λ i , T gas ) is the transmittance of the high-temperature gas;

[0042] ε gas (λ i , T gas ) is the emissivity of the high-temperature gas;

[0043] L b (λ i , T b) is the radiant energy of the turbine blade in the aero-engine;

[0044] L(λ i ,T gas ) is the radiant energy of the high-temperature gas in the aero-engine;

[0045] T gas and T b are the temperature of the high-temperature gas and the temperature of the turbine blade respectively.

[0046] Preferably, the relationship between the transmittance and the emissivity of the high-temperature gas is: τ gas (λ i ,T gas )+ε gas (λ i ,T gas )=1;

[0047] Substituting it into formula (1-1), the radiant energy of the high-temperature gas is obtained as:

[0048] L total =τ gas (λ i ,T gas )L b (λ i ,T b )+[1-τ gas (λ i ,T gas )]L(λ i ,T gas ) (1-2)

[0049] Calculating the radiant energy of the high-temperature gas in the corresponding band according to formula (1-2) is:

[0050]

[0051] Select two narrow bands in the characteristic absorption band of the high-temperature gas and denote them as λ1 and λ2 respectively. Then the radiant energy of the high-temperature gas at the bands λ1 and λ2 is:

[0052]

[0053]

[0054]

[0055] In order to eliminate the fluctuation of the single-point value caused by the spectral noise in the measurement process, formula (1-4) is extended to the form of band integration. At this time, the radiant energy of the high-temperature gas received in the two bands λ1 and λ2 is:

[0056]

[0057] Compared with the existing technology, the present invention realizes real-time measurement of the internal gas radiation spectrum of an aeroengine through a gas spectrum acquisition device, proposes a passive spectral dual-wavelength constraint method, and according to the constraint condition that the measured radiation spectrum and the gas temperature remain unchanged during the measurement process.

[0058] The present invention does not require an active light source, relies on measuring the internal radiation spectrum distribution during the operation of the aeroengine to invert the gas temperature, and has advantages such as good measurement real-time performance, simple modification of existing engines, and simple measurement devices compared with traditional gas temperature measurement methods, reducing the application cost and improving the stability of the measurement system. Description of the Drawings

[0059] Figure 1 is a schematic structural diagram of a high-temperature gas spectrum acquisition device for an aeroengine provided according to an embodiment of the present invention.

[0060] Figure 2 is a schematic flow diagram of a high-temperature gas temperature inversion method for an aeroengine provided according to an embodiment of the present invention.

[0061] Figure 3 is a program block diagram of a high-temperature gas temperature inversion method for an aeroengine provided according to an embodiment of the present invention.

[0062] The reference numerals therein include: turbine blade 1, measured high-temperature gas 2, optical probe system 3, condenser lens group 4, field stop 5, collimator lens group 6, optical fiber 7, fiber optic spectrometer 8, and data processing system 9. Detailed Embodiments

[0063] Hereinafter, embodiments of the present invention will be described with reference to the drawings. In the following description, the same modules are denoted by the same reference numerals. In the case of the same reference numerals, their names and functions are also the same. Therefore, their detailed descriptions will not be repeated.

[0064] In order to make the objectives, technical solutions, and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the drawings and specific embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and do not constitute a limitation to the present invention.

[0065] Figure 1 shows a schematic structural diagram of a high-temperature gas spectrum acquisition device for an aeroengine provided according to an embodiment of the present invention.

[0066] As Figure 1As shown, the high-temperature gas concentration and temperature measurement system for an aero-engine includes a turbine blade 1, a measured high-temperature gas 2, an optical probe system 3, an optical fiber 7, an optical fiber spectrometer 8, and a data processing system 9.

[0067] Among them, the optical probe system 3 is a nickel-based superalloy housing and an optical sight tube containing a tubular optical component. A rotatable nickel metal mirror is installed at the top of the optical probe system 3, and the rotation of the mirror is controlled by pushing a push rod by a closed-loop stepper motor to change the measurement angle of the optical system.

[0068] The optical sight tube includes a condenser lens group 4, a field stop 5, and a collimator lens group 6.

[0069] Among them, the condenser lens group 4 focuses the radiation energy from the engine through a multi-lens combination, enabling the optical probe system 3 to obtain better energy collection effects;

[0070] The field stop 5 is used to determine the shape and focal distance of the measured high-temperature gas 2. The collimation optical system transmits the radiation energy collected by the mirror from the aero-engine to the optical fiber spectrometer 8 through the optical fiber 7. After passing through the grating spectroscopy system in the optical fiber spectrometer 8, it is finally received by the data processing system 9 to obtain the internal radiation spectrum of the aero-engine.

[0071] Figure 2 The flowchart of the method for inverting the temperature of high-temperature gas in an aero-engine provided by an embodiment of the present invention is shown.

[0072] Figure 3 The program block diagram of the method for inverting the temperature of high-temperature gas in an aero-engine provided by an embodiment of the present invention is shown.

[0073] As Figure 2 and Figure 3 shown, the method for inverting the temperature of high-temperature gas in an aero-engine provided by an embodiment of the present invention includes the following steps:

[0074] S1. Obtain the radiation spectrum of the high-temperature gas through the high-temperature gas spectrum acquisition device of the aero-engine;

[0075] S2. Determine the characteristic absorption peak band by analyzing the spectral characteristics of the high-temperature gas;

[0076] S3. Establish a two-band gas radiation transfer model based on the characteristic absorption peak band of the high-temperature gas;

[0077] At the characteristic absorption band of the high-temperature gas spectrum, the radiation energy received by the high-temperature gas spectrum acquisition device of the aero-engine can be expressed as:

[0078] L total = τ gas (λ i ,Tgas )L b (λ i ,T b ) + [1 - τ gas (λ i ,T gas )]L(λ i ,T gas ) (1 - 1)

[0079] where

[0080] τ gas (λ i ,T gas ) is the transmittance of the high - temperature gas;

[0081] ε gas (λ i ,T gas ) is the emissivity of the high - temperature gas;

[0082] L b (λ i ,T b ) is the radiation energy of the turbine blade in the aero - engine;

[0083] L(λ i ,T gas ) is the radiation energy of the high - temperature gas in the aero - engine;

[0084] T gas and T b are the temperature of the high - temperature gas and the temperature of the turbine blade respectively.

[0085] The relationship between the transmittance and emissivity of the high - temperature gas is: τ gas (λ i ,T gas ) + ε gas (λ,iT ga ) s = 1.

[0086] Formula (1 - 1) can be expressed as:

[0087] L total = τ gas (λ i ,T gas )L b (λ i ,T b ) + [1 - τ gas (λ i ,T gas )]L(λ i ,T gas ) (1 - 2)

[0088] Since the response band of the spectrometer in the present invention is in the 900 - 2500 nm band and there is no high - intensity (characteristic) absorption peak of high - temperature gas, at the characteristic absorption peak of high - temperature gas, the energy received by the optical system includes both the radiation energy from the background and the radiation energy from the gas.

[0089] Based on the above, the present invention proposes a dual - band absorption coefficient ratio calibration algorithm to calculate the gas temperature, and calculates the gas temperature by calibrating the absorption coefficient ratio at the characteristic absorption band of the gas to be measured.

[0090] Taking carbon dioxide gas as an example for illustration, according to the gas spectral radiation transfer model to calculate the gas temperature, within the gas characteristic band range, two adjacent bands are respectively denoted as λ1 and λ2, and the radiation energy received by the spectrometer at the corresponding bands is calculated according to Equation (1 - 2):

[0091]

[0092] Select two adjacent narrow bands within the characteristic absorption band of high - temperature gas, denoted as λ1 and λ2 respectively, and the radiation energy received by the spectrometer at the corresponding bands is:

[0093]

[0094] In order to eliminate the fluctuation of single - point values caused by spectral noise during the measurement process, it is extended to the form of band integration. At this time, the radiation energy received from high - temperature gas within the two bands is:

[0095]

[0096] Among them,

[0097] λ1 and λ2 are respectively two narrow bands within the characteristic absorption band of high - temperature gas;

[0098] τ gas (λ i ,T gas ) is the transmittance of high - temperature gas;

[0099] T gas and T b are respectively the temperature of high - temperature gas and the temperature of the turbine blade;

[0100] E total (λ1) is the total radiation energy received by the spectral measurement device at the characteristic band λ1;

[0101] E total (λ2) is the total radiation energy received by the spectral measurement device at the characteristic band λ2;

[0102] Δλ1 is the integration bandwidth at the characteristic band λ1;

[0103] Δλ2 is the integration bandwidth at the characteristic wavelength λ2;

[0104] c1 = 3.7418×10 -16 W·m 2 is the first radiation constant;

[0105] c2 = 1.4388×10 -2 m·K is the second radiation constant.

[0106] Combining with Planck's formula, the high-temperature gas temperature corresponding to each band can be obtained:

[0107]

[0108] S4. Construct a deviation function for solving the high-temperature gas temperature and determine the constraint conditions;

[0109] From the band distribution with strong absorption at the absorption peak position of the gas spectrum, select λ1 = 1.95μm, λ2 = 2.05μm, Δλ = 50μm. The temperatures of the high-temperature gas inverted by the two bands at the same moment should be the same. Therefore, the temperature deviation of the high-temperature gas calculated by the two bands should approach zero. Thus, the temperature-solving deviation function is constructed as:

[0110]

[0111] The deviation function contains three unknowns, and the non-linear optimization method is used to solve the gas temperature.

[0112] Solving the gas temperature is transformed into a non-linear constrained optimization problem. Among them, according to the distribution of the gas concentration in the engine and the optical path of 6 cm, and the transmittance distribution law of the high-temperature gas near the absorption peak position, the transmittance of the high-temperature gas layer is higher than 0.9, and the difference in transmittance between the two bands should be less than 0.1. Therefore, the comprehensive above measurement conditions, the constraint conditions are:

[0113]

[0114] S5. Solve the deviation function within the constraint conditions through the Levenberg-Marquard algorithm to obtain N deviation function values.

[0115] S6. Take the temperature T gas corresponding to the minimum deviation function value among the N deviation function values as the actual temperature value of the high-temperature gas.

[0116] Optimize the minimum value of the deviation function according to the constraint conditions. When the constraint conditions in the above formula are satisfied, it is the temperature of the high-temperature gas at this time.

[0117] Although the embodiments of the present invention have been shown and described above, it can be understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those of ordinary skill in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present invention.

[0118] The specific implementation manners of the present invention above do not constitute a limitation to the protection scope of the present invention. Any other corresponding changes and deformations made according to the technical concept of the present invention shall be included in the protection scope of the claims of the present invention.

Claims

1. An inversion method for the high-temperature gas temperature of an aero-engine, which is implemented based on a high-temperature gas concentration and temperature measurement system of the aero-engine, is characterized in that, The following steps are involved: S1. Obtaining the radiation spectrum of high-temperature fuel gas through an aircraft engine high-temperature fuel gas spectrum acquisition device; S2. Analyzing the spectral characteristics of the high-temperature fuel gas to determine the characteristic absorption peak band; S3. Establishing a dual-band gas radiation transmission model according to the characteristic absorption peak band of the high-temperature gas: in, λ1 and λ2 are two narrow bands within the characteristic absorption band of the high-temperature fuel gas; τ gas (λ i ,T gas ) is the transmittance of the high-temperature gas; T gas and T b are the temperature of the high-temperature gas and the temperature of the turbine blade, respectively; E total (λ1) is the total radiation energy received by the spectral measurement device at the characteristic wavelength band λ1; E total (λ2) is the total radiant energy received by the spectral measurement device at the characteristic wavelength band λ2; Δλ1 is the integrated bandwidth at the characteristic band λ1; Δλ2 is the integrated bandwidth at the characteristic band λ2; c1 = 3.7418×10 -16 W·m 2 is the first radiation constant; c2 = 1.4388×10 -2 m·K is the second radiation constant; S4, constructing the high-temperature fuel gas temperature solution deviation function and determining constraint conditions; The deviation function is: The constraints of the deviation function are: S5. Solving the deviation function within the constraints by using the Levenberg-Marquard algorithm to obtain N deviation function values; S6. Take the temperature T corresponding to the minimum deviation function value among the N deviation function values gas as the actual temperature value of the high-temperature gas.

2. The method for inverse calculation of high-temperature gas temperature of an aero-engine according to claim 1, wherein In step S3: At the characteristic absorption band of the high-temperature gas spectrum, the radiation energy received by the aircraft engine high-temperature gas spectrum acquisition device is: L total = τ gas (λ i , T gas )L b (λ i , T b ) + [1 - τ gas (λ i , T gas )]L(λ i , T gas ) (1 - 1) in, τ gas (λ i ,T gas ) is the transmittance of the high-temperature gas; ε gas (λ i ,T gas ) is the emissivity of the high-temperature gas; L b (λ i ,T b ) is the radiant energy of the turbine blade in the aero-engine; L(λ i ,T gas ) is the radiant energy of the high-temperature gas in the aero-engine; T gas and T b are respectively the temperature of the high-temperature gas and the temperature of the turbine blade.

3. The method for inverting the high-temperature gas temperature of an aero-engine according to claim 2, wherein The relationship between the transmittance and emissivity of the high-temperature gas is: τ gas (λ i ,T gas )+ε gas (λ i ,T gas ) = 1; Substituting it into the formula (1-1), the radiation energy of the high-temperature fuel gas is obtained as follows: L total = τ gas (λ i , T gas )L b (λ i , T b ) + [1 - τ gas (λ i , T gas )]L(λ i , T gas ) (1 - 2) According to the formula (1-2), the radiation energy of high-temperature gas in the corresponding band is calculated as: Select two narrow bands in the characteristic absorption band of the high-temperature gas and record them as λ1 and λ2 respectively. Then the radiation energy of the high-temperature gas at the bands λ1 and λ2 is: In order to eliminate the fluctuation of single-point value caused by spectral noise during the measurement process, the formula (1-4) is expanded into the form of band integration. At this time, the energy radiated by high-temperature gas in the two bands λ1 and λ2 is: The formula (1-5) is combined with the Planck formula to obtain the high-temperature gas temperature corresponding to each band:

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