CN113482960B - Method for judging surge of aviation gas turbine engine - Google Patents
Method for judging surge of aviation gas turbine engine Download PDFInfo
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/001—Testing thereof; Determination or simulation of flow characteristics; Stall or surge detection, e.g. condition monitoring
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Abstract
The application belongs to the field of aviation gas turbine engines, and particularly relates to a surge judgment method of an aviation gas turbine engine. The method comprises the following steps: step one, judging whether the engine is in a non-stop state, if so, carrying out the next step; step two, acquiring a high-voltage conversion rotating speed n2r, and judging whether the high-voltage conversion rotating speed n2r is greater than a first threshold value, if so, performing the next step; step three, obtaining a fan surge signal U1; judging whether the fan surge signal U1 is greater than a second threshold value U1max, if so, carrying out the next step; and/or obtaining a compressor surge signal U2; judging whether the surge signal U2 of the compressor is greater than a third threshold value U2max, if so, carrying out the next step; and step four, judging whether the engine is restarted, and if not, defining that the engine surging occurs. The method and the device can reduce the risk of surge, improve the pneumatic stability and reliability of the engine, and ensure the flight safety.
Description
Technical Field
The application belongs to the field of aviation gas turbine engines, and particularly relates to a surge judgment method of an aviation gas turbine engine.
Background
Surge is an unstable operation phenomenon of the compression components (fans, compressors) of an aircraft engine. When surging occurs, parameters such as flow and pressure of a fan and an air compressor oscillate with low frequency and high amplitude along with time, and backflow phenomenon can occur in severe cases. This may lead to mechanical wear and the like which may cause adverse effects such as reduced engine performance, turbine overheating, increased vibration and stress of the rotor blades, and even damage to the structural integrity of the engine, directly threatening flight safety.
The aviation gas turbine engines of multiple models are researched in China, and usually a surge differential pressure annunciator is arranged behind a gas compressor, and surging is judged by acquiring the total back pressure differential behind the gas compressor and adding a hardware processing circuit. For a new research and development motivation, a large amount of attack and shutdown work is often required to be carried out to carry out adaptive improvement on a surge differential pressure sensor and a hardware processing circuit, such as sensor calibration and improvement, external pipeline blowing and improvement, hardware-in-loop simulation, hardware circuit improvement, surge-forcing test and the like, so as to obtain soft and hardware parameters matched with the pneumatic characteristics of a developed engine, and the technology cannot accurately judge the first-stage part of surge. Before the work of the attack is not completed, the engine has to withstand the risks and hazards posed by surge. Once surging occurs, the mechanical failure of the engine can be caused by light conditions, such as the aggravation of the abrasion of engine parts, the loss or the degradation of the performance function of the engine, and heavy conditions, such as the mechanical failure of the engine, the huge loss is brought, the model development or the equipment use is delayed, and the development and use cost is greatly increased. If surging can be quickly identified and the initial stage position of surging can be determined through less work, the potential adverse effect brought by surging is reduced, and the method has very important significance on the improvement design and the stability expansion design of an engine compression part, the pneumatic stability and the reliability of an engine and the flight safety.
Accordingly, a technical solution is desired to overcome or at least alleviate at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
The application aims to provide a method for judging the surge of an aviation gas turbine engine, so as to solve at least one problem existing in the prior art.
The technical scheme of the application is as follows:
an aviation gas turbine engine surge judgment method comprises the following steps:
step one, judging whether the engine is in a non-stop state, if so, carrying out the next step;
step two, acquiring a high-voltage conversion rotating speed n2r, and judging whether the high-voltage conversion rotating speed n2r is greater than a first threshold value, if so, performing the next step;
step three, obtaining a fan surge signal U1:
U1＝2(Pps202(j–n)–Pps202(j))/(Pps202(j–n)+Pps202(j))
wherein Pps202(j) is the static pressure pulsating pressure between the fan stages at the current moment, and Pps202(j-n) is the static pressure pulsating pressure between the fan stages at the previous moment;
judging whether the fan surge signal U1 is greater than a second threshold value U1max, if so, carrying out the next step; and/or the presence of a gas in the gas,
acquiring a compressor surge signal U2:
U2＝2(Pps253(j–n)–Pps253(j))/(Pps253(j–n)+Pps253(j))
wherein Pps253(j) is the static pressure pulsating pressure of the compressor interstage static pressure at the current moment, and Pps253(j-n) is the static pressure pulsating pressure of the compressor interstage static pressure at the previous moment;
judging whether the surge signal U2 of the compressor is greater than a third threshold value U2max, if so, carrying out the next step;
and step four, judging whether the engine is restarted, and if not, defining that the engine surging occurs.
In at least one embodiment of the present application, in step two, the high-pressure reduced rotation speed n2r is:
n2r＝N2/N2d*SQRT(288.15/T2)
wherein, N2 is the engine high pressure rotor speed, N2d is the engine high pressure rotor design speed, and T2 is the total fan inlet temperature.
In at least one embodiment of the present application, the first threshold is 35-40%.
In at least one embodiment of the present application, the first threshold is 40%.
In at least one embodiment of the present application, in step three, the fan inter-stage static pressure pulsation pressure Pps202 is processed by the first low pass filter, and the time interval between the current time and the previous time is 30 ms.
In at least one embodiment of the present application, the cut-off frequency of the first low-pass filter is 40 Hz.
In at least one embodiment of the present application, in step three, the compressor inter-stage static pressure pulsating pressure Pps253 is processed by a second low-pass filter, and the time interval between the current time and the previous time is 30 ms.
In at least one embodiment of the present application, the cut-off frequency of the second low-pass filter is 50 Hz.
The invention has at least the following beneficial technical effects:
the method for judging the surge of the aviation gas turbine engine can solve the technical problem that the surge and the first-stage surge part cannot be effectively judged in the working process of a newly-developed engine, reduce the risk of surge generation, provide support for the improvement design and stability expansion design of a newly-developed engine compression part, improve the aerodynamic stability and reliability of the engine and ensure the flight safety.
Drawings
FIG. 1 is a flow chart of a method for determining aircraft gas turbine engine surge in accordance with an embodiment of the present application.
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. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are a subset of the embodiments in the present application and not all embodiments in the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application and should not be construed as limiting the present application. All other embodiments obtained by a person of ordinary skill in the art based on the embodiments in the present application without making any creative effort belong to the protection scope of the present application. Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
In the description of the present application, it is to be understood that the terms "central," "longitudinal," "lateral," "front," "back," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in the orientations and positional relationships indicated in the drawings, which are based on the orientation or positional relationship shown in the drawings, and are used for convenience in describing the present application and for simplicity in description, but do not indicate or imply that the devices or elements referred to must have a particular orientation, be constructed in a particular orientation, and be operated, and therefore should not be construed as limiting the scope of the present application.
The present application is described in further detail below with reference to fig. 1.
When surging occurs, periodic pulsation can occur in flow passage pressure parameters of a fan and a gas compressor, the surging characteristics and the layout condition of measuring points of the general engine actual test are integrated, and meanwhile, in order to determine the initial surge stage position of the surging, the application provides a surging judgment method of an aviation gas turbine engine, which comprises the following steps:
step one, judging whether the engine is in a non-stop state, if so, carrying out the next step;
step two, acquiring a high-voltage conversion rotating speed n2r, and judging whether the high-voltage conversion rotating speed n2r is greater than a first threshold value, if so, carrying out the next step;
step three, obtaining a fan surge signal U1:
U1＝2(Pps202(j–n)–Pps202(j))/(Pps202(j–n)+Pps202(j))
wherein Pps202(j) is the static pressure pulsating pressure between the fan stages at the current moment, and Pps202(j-n) is the static pressure pulsating pressure between the fan stages at the previous moment;
judging whether the fan surge signal U1 is greater than a second threshold value U1max, if so, carrying out the next step; and/or the presence of a gas in the gas,
acquiring a compressor surge signal U2:
U2＝2(Pps253(j–n)–Pps253(j))/(Pps253(j–n)+Pps253(j))
wherein Pps253(j) is the static pressure pulsating pressure of the compressor interstage static pressure at the current moment, and Pps253(j-n) is the static pressure pulsating pressure of the compressor interstage static pressure at the previous moment;
judging whether the surge signal U2 of the compressor is greater than a third threshold value U2max, if so, carrying out the next step;
and step four, judging whether the engine is restarted, and if not, defining that the engine surging occurs.
According to the aviation gas turbine engine surge judgment method, the surge judgment is suitable for the working state of the engine, and the stop state does not need to be judged and treated, so that whether the engine is in the non-stop state or not is judged firstly. When the engine is in the non-stop state, it is determined whether or not the high-pressure converted rotation speed n2r is greater than a first threshold value in order to reduce the possibility of erroneous determination in a low state and a low pressure. The high-pressure converted rotation speed n2r is:
n2r＝N2/N2d*SQRT(288.15/T2)
wherein, N2 is the engine high pressure rotor speed, N2d is the engine high pressure rotor design speed, and T2 is the total fan inlet temperature.
In a preferred embodiment of the present application, the relative reduced rotation speed is generally greater than 35 to 40%, and the first threshold value is selected to be 35 to 40%. In this embodiment, the first threshold value is 40%.
The method for judging the surge of the aviation gas turbine engine comprises the steps of calculating a fan surge signal U1 through static pressure pulsating pressure Pps202 between fan stages, calculating a compressor surge signal U2 through static pressure pulsating pressure Pps253 between compressor stages, and taking at least one of the fan surge signal U1 and the compressor surge signal U2 which is larger than a corresponding threshold value as a judgment condition for the surge of the engine. In a preferred embodiment of the present application, the fan inter-stage static pressure pulsation pressure Pps202 is processed by a first low pass filter, and the time interval between the fan inter-stage static pressure pulsation pressure at the present moment and the measurement of the fan inter-stage static pressure pulsation pressure at the previous moment is selected to be 30ms, in this embodiment, the cut-off frequency of the first low pass filter is 40 Hz. The compressor interstage static pressure pulsating pressure Pps253 is processed through a second low-pass filter, the time interval between the compressor interstage static pressure pulsating pressure at the current moment and the time interval between the compressor interstage static pressure pulsating pressure at the previous moment in measurement is also selected to be 30ms, and in the embodiment, the cut-off frequency of the second low-pass filter is 50 Hz.
According to the aviation gas turbine engine surge judgment method, the second threshold value U1max and the third threshold value U2max are required to be obviously larger than the pulsation value in the normal acceleration and deceleration process, the transient process pulsation value can be obtained through a test, the pulsation value when the aircraft gas turbine engine enters surge is obtained through a surge approximation test, and the surge judgment given value is determined according to the obtained pulsation value.
In a preferred embodiment of the present application, the values of the second threshold value U1max and the third threshold value U2max may correspond to different values at different scaled rotation speeds. In the present embodiment, second threshold value U1max and third threshold value U2max of a certain engine are shown in tables 1 and 2.
TABLE 1
| n1r | 0.3 | 0.5 | 0.7 | 0.9 | 1.0 |
| U1max | 0.5 | 0.5 | 0.5 | 0.5 | 0.5 |
TABLE 2
| n2r25 | 0.7 | 0.8 | 0.9 | 1.0 |
| U2max | 0.4 | 0.4 | 0.4 | 0.4 |
Wherein,
n1r＝N1/N1d*SQRT(288.15/T2)
n1r is the converted rotating speed of the fan, N1 is the rotating speed of a low-pressure rotor of the engine, N1d is the designed rotating speed of the fan rotor of the engine, and T2 is the total temperature of an inlet of the fan;
n2r25＝N2/N2d*SQRT(T25d/T25)
n2r25 is the converted rotating speed of the compressor inlet, N2 is the rotating speed of the high-pressure rotor of the engine, N2d is the designed rotating speed of the compressor rotor of the engine, T25d is the total temperature of the compressor inlet at the design point, and T25 is the total temperature of the compressor inlet.
In the preferred embodiment of the present application, in order to accurately and timely reflect the variation of the pulsation parameters of the fan and the compressor, the acquisition frequencies of the sensors, namely the fan inter-stage static pressure pulsation pressure Pps202 and the compressor inter-stage static pressure pulsation pressure Pps253, should meet certain requirements, for example, in a certain engineering project, the acquisition frequency is 1 kHz.
The aviation gas turbine engine surge judgment method has the advantages that after the engine is judged to be restarted, whether the engine surges is finally determined, so that misjudgment caused by the fact that the instantaneous pulse value of restarting ignition exceeds the given value is avoided.
The method for judging the surge of the aviation gas turbine engine can judge the surge of the aviation gas turbine engine by measuring the parameters of the static pressure pulsation Pps202 of the fan interstage, the static pressure pulsation Pps253 of the compressor interstage, the total temperature T2 of a fan inlet, the total temperature T25 of the compressor inlet, the low-pressure rotating speed N1, the high-pressure rotating speed N2 and other fan and compressor interstage pulsation parameters, solves the technical problem that the surge and the first stage part of the surge cannot be effectively judged in the working process of a newly developed engine under the existing scheme, reduces the risk of the surge generation, provides support for the improvement design and the stability expansion design of the newly developed engine compression part, improves the aerodynamic stability and reliability of the engine and ensures the flight safety. Meanwhile, a new idea is provided for solving similar problems in the development of other models, so that the model development is smoothly carried out, and the model development cost is reduced. In addition, the application is simple to implement, easy to improve and wide in adaptability.
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 (8)
1. An aviation gas turbine engine surge judgment method is characterized by comprising the following steps:
step one, judging whether the engine is in a non-stop state, if so, carrying out the next step;
step two, acquiring a high-voltage conversion rotating speed n2r, and judging whether the high-voltage conversion rotating speed n2r is greater than a first threshold value, if so, performing the next step;
step three, obtaining a fan surge signal U1:
U1＝2(Pps202(j–n)–Pps202(j))/(Pps202(j–n)+Pps202(j))
wherein Pps202(j) is the static pressure pulsating pressure between the fan stages at the current moment, and Pps202(j-n) is the static pressure pulsating pressure between the fan stages at the previous moment;
judging whether the fan surge signal U1 is greater than a second threshold value U1max, if so, carrying out the next step; and/or the presence of a gas in the gas,
acquiring a compressor surge signal U2:
U2＝2(Pps253(j–n)–Pps253(j))/(Pps253(j–n)+Pps253(j))
wherein Pps253(j) is the static pressure pulsating pressure of the compressor interstage static pressure at the current moment, and Pps253(j-n) is the static pressure pulsating pressure of the compressor interstage static pressure at the previous moment;
judging whether the surge signal U2 of the compressor is greater than a third threshold value U2max, if so, carrying out the next step;
and step four, judging whether the engine is restarted, and if not, defining that the engine surging occurs.
2. The aircraft gas turbine engine surge determination method according to claim 1, wherein in step two, the high-pressure reduced rotation speed n2r is:
n2r＝N2/N2d*SQRT(288.15/T2)
wherein, N2 is the engine high pressure rotor speed, N2d is the engine high pressure rotor design speed, and T2 is the total fan inlet temperature.
3. The aircraft gas turbine engine surge determination method of claim 2, wherein the first threshold value is 35-40%.
4. The aircraft gas turbine engine surge determination method of claim 3, wherein the first threshold value is 40%.
5. The aircraft gas turbine engine surge judging method as claimed in claim 1, wherein in the third step, the static pressure pulsation pressure Pps202 between the fan stages is processed by a first low pass filter, and the time interval between the current time and the previous time is 30 ms.
6. The aircraft gas turbine engine surge determination method of claim 5, wherein a cutoff frequency of the first low pass filter is 40 Hz.
7. The aircraft gas turbine engine surge judging method as claimed in claim 1, wherein in the third step, the compressor inter-stage stator static pressure pulsating pressure Pps253 is processed by a second low-pass filter, and the time interval between the current time and the previous time is 30 ms.
8. The aircraft gas turbine engine surge determination method of claim 7, wherein a cutoff frequency of the second low pass filter is 50 Hz.
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| CN113931866B (en) * | 2021-10-27 | 2023-06-20 | 中国航发沈阳发动机研究所 | Pneumatic instability identification method for aero-engine compressor |
| CN114183251A (en) * | 2021-11-08 | 2022-03-15 | 陕西千山航空电子有限责任公司 | Surge protection control method for aircraft engine |
| CN114112176B (en) * | 2021-11-10 | 2023-09-22 | 中国航发沈阳发动机研究所 | Design method of external pipeline connected with surge pressure difference or pressure sensor |
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