CN115688450A - Method for monitoring temperature in front of turbine in over-temperature test of whole aircraft engine - Google Patents
Method for monitoring temperature in front of turbine in over-temperature test of whole aircraft engine Download PDFInfo
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- CN115688450A CN115688450A CN202211391581.7A CN202211391581A CN115688450A CN 115688450 A CN115688450 A CN 115688450A CN 202211391581 A CN202211391581 A CN 202211391581A CN 115688450 A CN115688450 A CN 115688450A
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Abstract
The application provides a method for monitoring the front temperature of a turbine in the overtemperature test of the whole aircraft engine, which comprises the following steps: obtaining the relation between the inlet temperature of a first-stage turbine rotor of a high-pressure turbine and the outlet temperature of the high-pressure turbine and the relation between the inlet temperature of a low-pressure turbine of the low-pressure turbine and the rear exhaust temperature of the turbine according to a high-pressure turbine work balance equation and a low-pressure turbine work balance equation, wherein the outlet temperature of the high-pressure turbine rotor is equal to the inlet temperature of the low-pressure turbine, and the inlet temperature of the first-stage turbine rotor and the rear exhaust temperature of the turbine are compared to obtain a proportionality coefficient; and establishing a relational expression containing the turbine rear exhaust temperature, the first-stage turbine rotor inlet temperature and the proportional coefficient in the overtemperature state, measuring to obtain the turbine rear exhaust temperature in the engine overtemperature state, and obtaining the first-stage turbine rotor inlet temperature at the lower end of the overtemperature state through the relational expression and the measured value of the turbine rear exhaust temperature in the overtemperature state, thereby realizing the monitoring of the first-stage turbine rotor inlet temperature in the overtemperature state.
Description
Technical Field
The application belongs to the technical field of aero-engines, and particularly relates to a method for monitoring the front temperature of a turbine in an over-temperature test of a complete machine of an aero-engine.
Background
The temperature of the hot end part exceeds the highest allowable gas temperature possibly caused by control system failure or other reasons in the use process of the engine, so that the novel aviation turbojet/turbofan engine needs to carry out an overtemperature test in a specified test before flight and a shaped structural test to verify the structural integrity of an engine rotating member and the structural strength of the hot end part, a test basis is provided for the continued use of the aero-engine after overtemperature, and the use safety of an airplane equipped with a single engine can be greatly improved by the method.
The overtemperature test is a test item specified in relevant standards which states that "the overtemperature test should be operated at least 45 ℃ above the steady state maximum allowable gas temperature at the inlet of the first stage turbine rotor and at not less than the steady state maximum allowable rotational speed for at least 5 minutes, the dimensions of parts and components after the test are within the allowable limits, and there is no sign of imminent failure, i.e. the test is considered to be satisfactorily completed".
Although there are specific index requirements for the over-temperature test in the relevant standards, no explicit provision is made for a method of monitoring the maximum allowable inlet steady state gas temperature of the first stage turbine rotor. Along with the gradual improvement of the performance of modern aircraft engines, the temperature and the pressure of the inlet temperature measuring position of the first-stage turbine rotor are very high when the engines are in an overtemperature state, and the temperature of the section cannot be directly measured due to the limitation of structural space and the like, so that the development of an overtemperature examination test of the whole aircraft is not facilitated.
Disclosure of Invention
The application aims to provide a method for monitoring the front temperature of a turbine in the overtemperature test of the whole aircraft engine, so as to solve or alleviate at least one problem in the background art.
The technical scheme of the application is as follows: a method for monitoring the front temperature of a turbine in an aircraft engine complete machine over-temperature test comprises the following steps:
obtaining the relation between the inlet temperature of a first-stage turbine rotor of a high-pressure turbine and the outlet temperature of the high-pressure turbine and the relation between the inlet temperature of a low-pressure turbine of the low-pressure turbine and the rear exhaust temperature of the turbine according to a high-pressure turbine work balance equation and a low-pressure turbine work balance equation, wherein the outlet temperature of the high-pressure turbine rotor is equal to the inlet temperature of the low-pressure turbine, and the inlet temperature of the first-stage turbine rotor and the rear exhaust temperature of the turbine are compared to obtain a proportionality coefficient;
and establishing a relational expression containing the turbine rear exhaust temperature, the first-stage turbine rotor inlet temperature and the proportional coefficient in the overtemperature state, measuring to obtain the turbine rear exhaust temperature in the engine overtemperature state, and obtaining the first-stage turbine rotor inlet temperature at the lower end of the overtemperature state through the relational expression and the measured value of the turbine rear exhaust temperature in the overtemperature state, thereby realizing the monitoring of the first-stage turbine rotor inlet temperature in the overtemperature state.
Further, the work balance equation of the high-pressure turbine is as follows:
in the formula (I), the compound is shown in the specification,
is the high pressure turbine average specific heat;
T 41 is the first stage turbine rotor inlet temperature;
T 45 is the high pressure turbine rotor inlet temperature;
κ 4 is the specific heat ratio of the high-pressure turbine;
π TH a high pressure turbo expansion ratio;
η TH is the high pressure turbine efficiency.
Further, the inlet temperature of the first stage turbine rotor of the high pressure turbine and the outlet temperature of the high pressure turbine rotorThe relationship of (1) is:
in the formula, y 1 As a coefficient related to the high-pressure turbo expansion ratio,
further, the low-pressure turbine work balance equation is as follows:
in the formula (I), the compound is shown in the specification,
is the low pressure turbine average specific heat;
T 6 is the post-turbine exhaust temperature;
κ 45 is the low pressure turbine average specific heat;
π TL a low pressure turbo expansion ratio;
η TL is the low pressure turbine efficiency.
Further, the relationship between the low-pressure turbine inlet temperature of the low-pressure turbine and the post-turbine exhaust temperature is as follows:
in the formula, y 2 As a coefficient related to the low-pressure turbo expansion ratio,
further, the proportionality coefficient
Further, the relation formula of the turbine rear exhaust temperature, the first stage turbine rotor inlet temperature and the proportionality coefficient under the overtemperature condition is：T 6.mb ＝T 41.mb /k CW.A8
In the formula, T 6.mb The target measurement value of the turbine rear exhaust temperature in the overtemperature state is obtained;
T 41.mb the target monitoring value is the inlet temperature target of the first stage turbine rotor in an overtemperature state;
k cw.A8 for corresponding nozzle area A 8 Lower turbine rear exhaust temperature T 6 And the first stage turbine rotor inlet temperature T 41 Corresponding coefficients of (c).
The indirect monitoring method of the inlet temperature of the first-stage turbine rotor in the overtemperature state is provided aiming at the problem that the inlet temperature T41 of the first-stage turbine rotor cannot be directly measured in the overtemperature test, thereby realizing the representation of the inlet temperature of the first-stage turbine rotor by monitoring the exhaust temperature behind the turbine, avoiding the damage to the original structure of the engine caused by the fact that the inlet temperature of the first-stage turbine rotor in the overtemperature state is measured by mounting and measuring the sensitive part, reducing the extra test risk of the heat end part examined by the engine and the influence on the test result, and being beneficial to the smooth completion of the overtemperature examination test of the whole engine.
Drawings
In order to more clearly illustrate the technical solutions provided by the present application, the following briefly introduces the accompanying drawings. It is to be understood that the drawings described below are merely exemplary of some embodiments of the application.
Fig. 1 is a schematic view of a monitoring method of the present application.
FIG. 2 is a schematic cross-sectional view of a turbofan engine according to an embodiment of the present application.
FIG. 3 shows the T values at different nozzle opening areas in the present application 41 /T 6 At atmospheric temperature T 1 Graph of the relationship of (c).
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application.
As shown in FIG. 1, the indirect monitoring method for the turbine front temperature of the aircraft engine overtemperature test provided by the application comprises the following steps:
s1, obtaining the inlet temperature T of a first-stage turbine rotor of the high-pressure turbine according to a work balance equation of the high-pressure turbine and a work balance equation of the low-pressure turbine 41 And the high-pressure turbine rotor outlet temperature T 45 And a low-pressure turbine inlet temperature of the low-pressure turbine (i.e., a high-pressure turbine rotor outlet temperature T) 45 ) And turbine rear exhaust temperature T 6 Relating the first stage turbine rotor inlet temperature T to 41 And turbine rear exhaust temperature T 6 And (5) making a ratio to obtain a proportionality coefficient of the two.
As shown in fig. 2, which is a schematic diagram of component composition and cross-sectional identification of a typical aircraft engine, the aircraft engine sequentially includes a fan 10, a compressor 20, a combustion chamber 30, a high-pressure turbine 40, a low-pressure turbine 50, and a nozzle 60 along an airflow direction, the front end of the fan 10 is an engine inlet 1 and a fan inlet 2, the front end of the high-pressure turbine 40 is a high-pressure turbine inlet 41, the rear end of the high-pressure turbine 40 is a high-pressure turbine outlet 45, the high-pressure turbine outlet 45 is an inlet of the low-pressure turbine 50, the rear end of the low-pressure turbine 50 is a turbine rear exhaust outlet 6, and the tail of the nozzle 60 is an engine nozzle cross-section 8.
When engine inlet temperature T 1 For a certain time, the work balance of the high-pressure turbine can be obtained:
formula (1) is simplified to obtain:
in the formula (I), the compound is shown in the specification,
is the high pressure turbine average specific heat;
κ 4 is the specific heat ratio of the high-pressure turbine;
π TH a high pressure turbo expansion ratio;
η TH high pressure turbine efficiency;
y 1 as a coefficient related to the high-pressure turbo expansion ratio,
the work balance of the low-pressure turbine can be obtained:
further simplification of formula (3) gives:
in the formula (I), the compound is shown in the specification,
is the low pressure turbine average specific heat;
κ 45 is the low pressure turbine specific heat ratio;
π TL a low pressure turbo expansion ratio;
η TL low pressure turbine efficiency;
y 2 as a coefficient related to the low-pressure turbo expansion ratio,
the formula (4) is compared with the formula (2) to obtain a proportionality coefficient
As can be seen from equation (5), the proportionality coefficient k CW Related to the pressure drop ratio of high-pressure turbine and low-pressure turbine, and the nozzle area A of engine 8 Directly affects the change of the turbine pressure drop ratio of the engine, therefore when the nozzle area A of the engine is changed 8 A timing, coefficient k CW Is a certain value.
S2, establishing the exhaust temperature T after the turbine in the overtemperature state 6.mb And the first stage turbine rotor inlet temperature T 41.mb Measured to obtain the turbine in the over-temperature state of the engineRear exhaust temperature T 6.mb And obtaining the inlet temperature T of the first stage turbine rotor through the relational expression 41.mb Thereby realizing the inlet temperature T of the first stage turbine rotor 41.mb Monitoring of turbine rear exhaust temperature T in over-temperature conditions 6.mb And the first stage turbine rotor inlet temperature T 41.mb The relation of (A) is as follows: t is a unit of 6.mb ＝T 41.mb /k CW.A8 ； (6)
In the formula, T 6.mb The target measurement value of the exhaust temperature after the turbine is in an overtemperature state;
T 41.mb the target monitoring value is the inlet temperature target of the first stage turbine rotor in an overtemperature state;
k cw.A8 corresponding nozzle area A 8 Lower turbine rear exhaust temperature T 6 And the first stage turbine rotor inlet temperature T 41 The corresponding coefficient of (a).
FIG. 3 shows different engine throat areas A provided in the embodiment of the present application 8 T of 41 /T 6 And a monitoring value capable of accurately representing the inlet temperature of the first stage turbine rotor can be obtained through the relation curve of the ratio and the inlet temperature T1 of the engine and the measurement of the exhaust temperature behind the turbine.
The indirect monitoring method of the inlet temperature of the first-stage turbine rotor under the overtemperature condition is provided aiming at the problem that the inlet temperature T41 of the first-stage turbine rotor cannot be directly measured during the overtemperature test, thereby realizing the monitoring of the inlet temperature of the first-stage turbine rotor by exhaust temperature after the turbine, avoiding the damage to the original structure of the engine caused by the installation and measurement of the sensing part due to the fact that the inlet temperature of the first-stage turbine rotor under the overtemperature condition is directly measured, reducing the extra test risk of the examined hot end part of the engine and the influence on the test result, and being beneficial to the smooth completion of the overtemperature examination test of the whole engine.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (7)
1. A method for monitoring the front temperature of a turbine in an aircraft engine complete machine over-temperature test is characterized by comprising the following steps:
obtaining the relation between the inlet temperature of a first-stage turbine rotor of the high-pressure turbine and the outlet temperature of the high-pressure turbine rotor and the relation between the inlet temperature of a low-pressure turbine of the low-pressure turbine and the exhaust temperature after the turbine according to a high-pressure turbine work-doing balance equation and a low-pressure turbine work-doing balance equation, wherein the outlet temperature of the high-pressure turbine rotor is equal to the inlet temperature of the low-pressure turbine, and the inlet temperature of the first-stage turbine rotor and the exhaust temperature after the turbine are compared to obtain a proportionality coefficient;
and establishing a relational expression containing the turbine rear exhaust temperature, the first-stage turbine rotor inlet temperature and the proportional coefficient in the overtemperature state, measuring to obtain the turbine rear exhaust temperature in the engine overtemperature state, and obtaining the first-stage turbine rotor inlet temperature at the lower end of the overtemperature state through the relational expression and the measured value of the turbine rear exhaust temperature in the overtemperature state, thereby realizing the monitoring of the first-stage turbine rotor inlet temperature in the overtemperature state.
2. The method for monitoring the front temperature of the turbine in the over-temperature test of the whole aircraft engine as claimed in claim 1, wherein the high-pressure turbine work balance equation is as follows:
in the formula (I), the compound is shown in the specification,
is the high pressure turbine average specific heat;
T 41 is the first stage turbine rotor inlet temperature;
T 45 is the high pressure turbine rotor inlet temperature;
κ 4 is the specific heat ratio of the high-pressure turbine;
π TH a high pressure turbo expansion ratio;
η TH is the high pressure turbine efficiency.
3. The method for monitoring the front temperature of the turbine in the whole aircraft engine over-temperature test of the aircraft engine as claimed in claim 2, wherein the relationship between the inlet temperature of the first stage turbine rotor of the high-pressure turbine and the outlet temperature of the high-pressure turbine rotor is as follows:
in the formula, y 1 As a coefficient related to the high-pressure turbo expansion ratio,
4. the method for monitoring the front temperature of the turbine in the over-temperature test of the whole aircraft engine as claimed in claim 3, wherein the low-pressure turbine work-doing balance equation is as follows:
in the formula (I), the compound is shown in the specification,
is the low pressure turbine average specific heat;
T 6 is the post-turbine exhaust temperature;
κ 45 is the low pressure turbine specific heat ratio;
π TL a low pressure turbo expansion ratio;
η TL is the low pressure turbine efficiency.
5. The method for monitoring the pre-turbine temperature in the over-temperature test of the whole aircraft engine as claimed in claim 4, wherein the inlet temperature and the turbine of the low-pressure turbine are respectively measured by a temperature sensor and a temperature sensorThe relationship of the exhaust gas temperature after the wheel is:
in the formula, y 2 As a coefficient related to the low-pressure turbo expansion ratio,
6. the method for monitoring the front temperature of the turbine in the over-temperature test of the whole aircraft engine as claimed in claim 5, wherein the proportionality coefficient
7. The method for monitoring the front temperature of the turbine in the over-temperature test of the whole aircraft engine as claimed in claim 6, wherein the relationship between the rear exhaust temperature of the turbine in the over-temperature state, the inlet temperature of the first stage turbine rotor and the proportionality coefficient is as follows: t is 6.mb ＝T 41.mb /k CW.A8
In the formula, T 6.mb The target measurement value of the exhaust temperature after the turbine is in an overtemperature state;
T 41.mb the target monitoring value is the inlet temperature target of the first stage turbine rotor in an overtemperature state;
k cw.A8 corresponding nozzle area A 8 Lower turbine rear exhaust temperature T 6 And the first stage turbine rotor inlet temperature T 41 The corresponding coefficient of (a).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211391581.7A CN115688450A (en) | 2022-11-07 | 2022-11-07 | Method for monitoring temperature in front of turbine in over-temperature test of whole aircraft engine |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202211391581.7A CN115688450A (en) | 2022-11-07 | 2022-11-07 | Method for monitoring temperature in front of turbine in over-temperature test of whole aircraft engine |
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