Method for controlling stable margin of compression component under whole machine pressure distortion condition

Abstract

The application belongs to the technical field of aero-engines, and particularly relates to a method for controlling the stability margin of a compression component under the condition of whole-machine pressure distortion, comprising the following steps: S1: calculating an inlet distortion index prediction value W; then calculating a fan pressure ratio correction amount △FPR MAX and a compressor pressure ratio correction amount △CPR MAX according to the inlet distortion index W; S2: correcting a current working state compressor pressure ratio limit value CPR MAX by the compressor pressure ratio correction amount △CPR MAX to obtain a final value of the current working state compressor pressure ratio limit CPR LIMIT ; correcting a current working state fan pressure ratio limit value FPR MAX by the fan pressure ratio correction amount △FPR MAX to obtain a final value of the current working state fan pressure ratio limit FPR LIMIT ; S3: when the compressor pressure ratio CPR is greater than CPR LIMIT , reducing the main combustor fuel supply amount to ensure that the compressor pressure ratio is not greater than CPR LIMIT ; when the fan pressure ratio FPR is greater than FPR LIMIT , enlarging the tail nozzle throat area to ensure that the fan pressure ratio is not greater than FPR LIMIT .

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aero-engines, and particularly relates to a method for controlling the stable margin of a compression component under the condition of whole-machine pressure distortion. BACKGROUND

[0002] When a gas turbine engine is working, fuel is combusted in a main combustion chamber component to convert chemical energy into heat energy, which is converted into mechanical energy (or kinetic energy) by a turbomachinery (or a nozzle). In order to ensure that fuel can be reliably and effectively combusted in the main combustion chamber, the air medium needs to be warmed and pressurized, and the engine compression components (low-pressure compressor, high-pressure compressor, etc.) must work stably and reliably.

[0003] When an aircraft is maneuvered at a large angle of attack, pressure distortion occurs at the engine inlet. As the inlet distortion intensity increases, the stable working boundary of the compression component decreases. When the actual working point is higher than the stable working boundary point, the compression component surges, the energy conversion process is interrupted, and the engine cannot work normally, which affects the combat effectiveness of the aircraft. According to the search of public technical documents, there is currently no method for real-time control of the stable margin of the compression component under the condition of whole-machine pressure distortion.

[0004] Since there is currently no method for real-time control of the stable margin of the compression component under the condition of whole-machine pressure distortion, the stall margin of the engine compression component is indirectly controlled mainly by measuring the measurable parameters (fan speed, engine pressure ratio) reflecting the working state of the engine by sensors. This indirect control mode is not accurate for the stall margin of the engine. When the engine is designed, the fan stall margin must be maintained at 15% to 20%, and the compressor stall margin is about 30%. The method has the following disadvantages:

[0005] 1) It is impossible to judge the downward displacement of the stable working boundary of the compression component, and the main engine acceleration fuel flow and the tail nozzle throat area cannot be adjusted, and problems such as stall and surge may still occur; 2) The reserved stall margin is large, which limits the performance of the engine; 3) Increases the workload of the pilot, affects the energy of the pilot, and is not conducive to the realization of the flight mission. SUMMARY

[0006] In order to solve the above problems, the application provides a method for controlling the stable margin of a compression component under the condition of whole-machine pressure distortion, which comprises:

[0007] Step S1: calculating the current working state compressor pressure ratio limit value CPR according to the compressor inlet equivalent flow W A25R and the compressor CPR limit line; MAX

[0008] calculating the current working state compressor pressure ratio limit value CPR according to the engine inlet equivalent flow W A2R ​and the fan FPR limit line, the current working state fan pressure ratio limit value FPR is calculated MAX ;

[0009] According to the engine inlet total pressure P t2 and the aircraft inlet total pressure P t0 , the aircraft inlet total pressure loss SIGINLET is calculated;

[0010] The engine inlet equivalent flow rate W A2R , the inlet total pressure loss SIGINLET, and the relationship of the inlet distortion index W are obtained;

[0011] According to the current state aircraft inlet total pressure loss SIGINLET, the engine inlet equivalent flow rate W A2R , and the above-mentioned relationship, the inlet distortion index prediction value W is calculated; and according to the compressor pressure ratio and the fan pressure ratio correction relationship curve under different inlet distortion index W conditions, the fan pressure ratio correction amount △FPR MAX and the compressor pressure ratio correction amount △CPR MAX are respectively calculated;

[0012] Step S2: The current working state compressor pressure ratio limit value CPR MAX is corrected by the compressor pressure ratio correction amount △CPR MAX to obtain the current working state compressor pressure ratio limit final value CPR LIMIT ; and the current working state fan pressure ratio limit value FPR MAX is corrected by the fan pressure ratio correction amount △FPR MAX to obtain the current working state fan pressure ratio limit final value FPR LIMIT ;

[0013] Step S3: When the engine is working:

[0014] When the compressor pressure ratio CPR is greater than CPR LIMIT , the main combustor fuel supply is reduced to ensure that the compressor pressure ratio is not greater than CPR LIMIT ;

[0015] When the fan pressure ratio FPR is greater than FPR LIMIT , the tail nozzle throat area is enlarged to ensure that the fan pressure ratio is not greater than FPR LIMIT .

[0016] Preferably, the method for obtaining the engine inlet equivalent flow rate W A2R in step S1 comprises:

[0017] The fan relative equivalent rotational speed N is obtained, and the fan inlet equivalent flow rate W is calculated based on the fan relative equivalent rotational speed N and the fan inlet total pressure PA2RJC The relationship calculation obtains the fan inlet conversion flow W A2RJC ;

[0018] Based on the influence of the fan rotor gap, the fan adjustable guide vane angle, and the Reynolds number, the fan inlet conversion flow basic value W A2RJC is corrected to obtain the engine inlet conversion flow W A2R .

[0019] Preferably, the method for obtaining the fan pressure ratio FPR in step S3 comprises:

[0020] Obtaining the low-pressure conversion speed and the fan bypass pressure ratio P t13 / P t2 ;

[0021] Obtaining the low-pressure conversion speed and the fan bypass pressure ratio P t13 / P t2 and the fan pressure ratio FPR relationship curve;

[0022] The calculation obtains the fan pressure ratio FPR.

[0023] Preferably, the method for obtaining the compressor pressure ratio CPR in step S3 comprises:

[0024] Obtaining the low-pressure conversion speed and the fan pressure ratio FPR and the fan temperature ratio T t21 / T t2 relationship curve, and linearly interpolating to obtain the fan temperature ratio T t21 / T t2 ;

[0025] According to the fan temperature ratio T t21 / T t2 , the fan inlet temperature measurement value T t2 , the fan outlet total temperature basic value T t21JC is calculated;

[0026] Based on the influence of the fan rotor gap, the Reynolds number, and other factors on the fan component efficiency, the fan outlet total temperature basic value T t21JC is corrected to obtain the fan outlet total temperature T t21 ; the correction formula is:

[0027] T t21 = T t21JC + DETATCL + DETATREU;

[0028] Wherein, DETATCL is the fan rotor gap correction value, and DETATREU is the Reynolds number correction value;

[0029] Based on the radial distribution difference between the total temperature at the fan outlet and the total temperature at the compressor inlet, the total temperature at the fan outlet T... t21 Correction yields compressor inlet temperature T t25 ;

[0030] Based on the relative physical rotation speed of high pressure The compressor inlet temperature T t25 The compressor's converted speed was calculated.

[0031] Obtain the pressure ratio P of the high-pressure compression component t3 / P t21 Obtain the compressor's converted speed High pressure compression component pressure ratio P t3 / P t21 Calculated flow rate W with the internal structure of the intermediate casing A23R Relationship curve;

[0032] The calculated inlet flow rate W of the intermediate casing is obtained. A23R ;

[0033] According to formula W A23 =W A23R ×(288.15 / T t21 ) 0.5 ×P t21 The inlet airflow W of the intermediate casing is calculated using / 101.325. A23 Among them, P t21 This refers to the fan outlet pressure.

[0034] According to the inlet airflow W inside the intermediate casing A23 The intermediate casing internal loss σ is obtained by calculating the relationship curve of the intermediate casing internal loss. NH ,

[0035] According to formula P t25 =P t21 ×σ NH The compressor inlet pressure P was calculated. t25 ;

[0036] Based on the compressor outlet pressure measurement value P t3 The compressor inlet pressure P t25 The compressor pressure ratio CPR is calculated using the formula CPR = P t3 / P t25 .

[0037] Preferably, the compressor inlet converted flow rate W A25R The calculation formula is:

[0038] W A25R =W A23 ×(Tt25 / 288.15) 0.5 x 101.325 / P t25 .

[0039] Preferably, the method for calculating the aircraft inlet total pressure loss SIGINLET in step S1 comprises:

[0040] Based on the W A2R ~ P t2 / P2 relationship curve, the inlet total pressure ratio P A2R / P2 value is calculated according to the inlet equivalent flow W t2 .

[0041] Based on the engine inlet static pressure measurement value P2, the inlet total pressure P t2 is calculated according to the total pressure ratio P t2 / P2 value.

[0042] The far front total pressure P t0 is calculated according to the ambient pressure P0 and the Mach number signal M given by the aircraft communication, and the calculation formula is: P t0 =P0 x (1+(k-1) x M 2 / 2) k / (k-1) .

[0043] The aircraft inlet total pressure loss SIGINLET is calculated according to the engine inlet total pressure P t2 and the aircraft inlet total pressure P t0 , and the calculation formula is: SIGINLET=P t2 / P t0 .

[0044] Preferably, the calculation formula of the final value of the current working state compressor pressure ratio limit CPR LIMIT is:

[0045] CPR LIMIT =CPR MAX -△CPR MAX .

[0046] The calculation formula of the final value of the current working state fan pressure ratio limit FPR LIMIT is:

[0047] FPR LIMIT =FPR MAX -△FPR MAX .

[0048] The application is based on engine model establishment method, import total pressure distortion intensity prediction and engine stability margin real-time control method to realize real-time control of compression component stability margin under whole machine distortion condition, effectively improve aircraft combat effectiveness, engine working safety and the like;

[0049] The engine model establishment method is as follows: according to theoretical calculation and test data, the engine model is constructed through W relationship curve, n 1r ~P t21 / P t2 ~T t21 / T t2 relationship curve and the like, and W A2R , W A25R , P 25 and the like parameters required for stability margin correction are obtained;

[0050] The import total pressure distortion intensity prediction is as follows: the aircraft inlet total pressure loss value is calculated based on engine import static pressure, the engine import comprehensive pressure distortion index is calculated by interpolation based on given low-pressure conversion speed, aircraft inlet total pressure loss and engine import comprehensive pressure distortion index relationship;

[0051] The engine stability margin real-time control method is as follows: according to the engine import total pressure distortion index W, the compression component pressure ratio upper limit value is corrected, and the compression component stability margin is controlled in real time by controlling main fuel flow and tail nozzle throat area to ensure that the compression component pressure ratio does not exceed the pressure ratio limit value. BRIEF DESCRIPTION OF DRAWINGS

[0052] Figure 1 is a method flow chart for controlling the compression component stability margin under whole machine pressure distortion condition according to a preferred embodiment of the application;

[0053] Figure 2 is a fan three common working line schematic diagram;

[0054] Figure 3 is a W A2R ~P t2 / P2 relationship curve schematic diagram;

[0055] Figure 4 is a W A25R ~CPRMAX relationship curve diagram;

[0056] Figure 5 is a W A2R ~FPRMAX relationship curve diagram;

[0057] Figure 6 Fan stability margin real-time correction diagram;

[0058] Figure 7 A compressor stability margin real-time correction map. DETAILED DESCRIPTION

[0059] In order to make the technical solutions of the present application and the advantages thereof clearer, the technical solutions of the present application will be further clearly and completely described in detail below in conjunction with the drawings. It should be understood that the specific embodiments described herein are only partial embodiments of the present application, and are only used to explain the present application, but not to limit the present application. It should be noted that, for the purpose of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design. In the case of no conflict, the embodiments in the present application and the technical features in the embodiments can be combined to obtain new embodiments.

[0060] In addition, unless otherwise defined, the technical terms or scientific terms used in the present application description should be the general meanings understood by the general technical personnel in the field to which the present application belongs. The words indicating the direction or position relationship such as "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer" and the like used in the present application description are only used to indicate the relative direction or position relationship, but not to imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation, and the relative position relationship thereof can also change accordingly when the absolute position of the described object changes, therefore it cannot be understood as a limitation on the present application. The "first", "second", "third" and the like used in the present application description are only for the purpose of description, and are used to distinguish different components, but cannot be understood as indicating or implying relative importance. The "one", "an" or "the" and the like used in the present application description should not be understood as an absolute limitation on the quantity, but should be understood as the existence of at least one. The "including" or "containing" and the like used in the present application description means that the elements or objects appearing before the word are encompassed by the elements or objects listed after the word and their equivalents, without excluding other elements or objects.

[0061] In addition, it should also be noted that, unless otherwise specified and limited, the "installation", "connection", "connection" and the like used in the present application description should be understood in a broad sense, for example, the connection can be a fixed connection, or a detachable connection, or an integral connection; can be a mechanical connection, or an electrical connection; can be directly connected, or indirectly connected through an intermediate medium, or the internal communication of two elements, and the person skilled in the art can understand the specific meaning of the present application according to the specific circumstances.

[0062] 1) Obtain the relative conversion speed of the fan by theoretical calculation or test method The relationship curve of the fan inlet conversion flow W A2RJC ;

[0063] 2) Consider the rotor clearance of the fan, the adjustable guide vane angle of the fan, the Reynolds number effect, and convert the basic value of the flow rate W of the fan inlet A2RJC Make corrections to obtain the engine inlet converted flow rate W A2R , and the calculation formula is W A2R = W A2RJC + DETAWAR_CL1+ DETAWAR_ALFA1+ DETAWAR_REU;

[0064] 3) Obtain the low-pressure converted speed N according to theoretical calculation or test method t13 Fan bypass pressure ratio (P t2 / P t21 ), fan pressure ratio (P t2 / P t21 ) relationship curve (see Figure 2 );

[0065] 4) Obtain the low-pressure converted speed N Fan pressure ratio (P t2 / P t21 ) and fan temperature ratio (T t2 / T LR ) relationship curve according to theoretical calculation or test method

[0066] Figure 2 The relationship curve corresponding to the three common working lines is shown in the figure: t21 / P t2 ~ T t21 / T t2 Upper boundary curve (TOP), reference boundary curve (MID), reference lower boundary curve (DOWN);

[0067] 5) Based on the actual low-pressure converted speed N Fan bypass pressure ratio (P t13 / P t2 ), the relationship curve provided in item 3), linear interpolation to obtain fan pressure ratio FPR (P t21 / P t2 );

[0068] 6) Based on the actual low-pressure converted speed N Fan pressure ratio FPR (P t21 / P t2 ), the relationship curve provided in item 4), linear interpolation to obtain fan temperature ratio (T t21 / T t2 );

[0069] ​7) According to item 6, fan temperature ratio (T) t21 / T t2 Fan inlet temperature measurement value T t2 The base value of the total temperature at the fan outlet, T, was calculated. t21JC ;

[0070] 8) Considering the impact of factors such as fan rotor clearance and Reynolds number on the efficiency of fan components, it is necessary to adjust the T given in 7). t21JC After correction, the total fan outlet temperature T is obtained. t21 The corrected formula is T. t21 =T t21JC +DETATCL+DETATREU, where DETATCL is the fan rotor clearance correction value and DETATREU is the Reynolds number correction value;

[0071] 9) Considering the difference in radial distribution between the total temperature at the fan outlet and the total temperature at the compressor inlet, it is necessary to adjust the T value given in clause 8). t21 The compressor inlet temperature T is obtained after correction. t25 ;

[0072] 10) Based on the relative physical rotational speed of the high pressure Item 9) Obtain T t25 The compressor's converted speed was calculated. The calculation formula is

[0073] 11) Obtain the compressor's converted speed through theoretical calculations or experimental methods. High pressure compression component pressure ratio (P) t3 / P t21 ) and the converted flow rate W of the intermediate casing A23R Relationship curve;

[0074] 12) Obtained from item 10) P t3 / P t21 The calculated inlet flow rate W of the intermediate casing is obtained from the relationship curve provided in item 11). A23R ;

[0075] 13) W obtained from item 12) A23R The P obtained from item 5) t21 The T obtained from item (8) t21 The inlet air flow rate W of the intermediate casing was calculated. A23 The calculation formula is W A23 =W A23R ×

[0076] (288.15 / T t21 ) 0.5 ×P t21 / 101.325;

[0077] 14) W A23R , the loss of mass content in the intermediate casing (W A23R ~ σ NH ) relationship curve to calculate the loss of mass content in the intermediate casing σ NH ;

[0078] 15) σ NH , the fan outlet pressure P t21 to calculate the compressor inlet pressure P t25 , the calculation formula is P t25 = P t21 × σ NH ;

[0079] 16) W A23 , T t25 , P t25 , W A25R , the calculation formula is W A25R =W A23 ×(T t25 / 288.15) 0.5 ×101.325 / P t25 ;

[0080] 17) Based on the measured value of the compressor outlet pressure P t3 , P t25 , the calculation formula is CPR=P t3 / P t25 ;

[0081] 18) According to the theoretical calculation or test method, the relationship curve of the engine inlet equivalent flow W A2R and the engine inlet total static pressure ratio P t2 / P2 is obtained, which is shown in Figure 3 ;

[0082] 19) Based on the W A2R ~ P t2 / P2 relationship curve, the inlet equivalent flow W A2R value obtained from item 1) is calculated to obtain the inlet total static pressure ratio P t2 / P2 value;

[0083] 20) Based on the measured value of the engine inlet static pressure P2 and the total static pressure ratio P t2 / P2 value obtained from item 19), the inlet total pressure P t2 is calculated;

[0084] 21) The total pressure P at the far front is calculated based on the ambient pressure P0 and Mach number signal M provided by the aircraft communication. t0 The calculation formula is: P t0 =P0×(1+(k-1)×M 2 / 2) k / (k-1) ;

[0085] 22) The total engine inlet pressure P obtained from item 20) t2 The total air intake pressure P of the aircraft obtained from item 21) t0 The total pressure loss SIGINLET of the aircraft air intake is calculated using the formula: SIGINLET = P t2 / P t0 .

[0086] b) Predicting engine inlet distortion intensity

[0087] 1) Obtain the converted flow rate W at the engine inlet based on theoretical calculations or experimental methods. A2R The relationship between total pressure loss in the intake manifold (SIGINLET) and the inlet distortion index (W);

[0088] 2) W obtained from the engine model in item a) A2R The predicted value W of the import distortion index is obtained by interpolating the relationship curve provided by SIGINLET and item 1).

[0089] c) Calculate the pressure ratio limit value of the compression component.

[0090] 1) Based on the characteristics of compressor components, considering the differences in component characteristics under whole-machine conditions and reserving a certain margin, determine the compressor pressure ratio (CPR) limiting curve W. A25R ~CPR MAX See diagram Figure 4 ;

[0091] 2) W obtained from the engine model in item a) A25R The W obtained from item 1) A25R ~CPR MAX The relationship curve provides the compressor pressure ratio limit (CPR) for the current operating state. MAX .

[0092] 3) Based on the characteristics of the fan components, considering the differences in component characteristics under the overall machine conditions and reserving a certain margin, determine the fan pressure ratio limiting curve W. A2R ~FPR MAX See diagram Figure 5 ;

[0093] 4) W obtained from the engine model in item a) A2R The W obtained from item 3) A2R ~FPR MAXRelationship curve, obtain the current operating state fan pressure ratio limit value FPR MAX .

[0094] d) Real-time correction of compression component stability margin

[0095] 1) Based on theoretical calculations or experimental methods, obtain the compressor pressure ratio and fan pressure ratio correction curves W ~ ΔCPR under different W conditions for the fan and compressor components. MAX W~△FPR MAX ;

[0096] 2) Based on the predicted value of engine inlet distortion intensity W obtained in item b) and the relationship curve obtained in item 1), the fan and compressor pressure ratio correction ΔCPR is calculated. MAX , △FPR MAX ;

[0097] 3) Based on the CPRMAX and FPRMAX obtained in item c), and the correction amount obtained in item 2), calculate the final value of the current state pressure ratio limit. The calculation formula is: CPR LIMIT =CPR MAX -△CPR MAX ,FPR LIMIT =FPR MAX -△FPR MAX .

[0098] 4) When the engine is running, the compressor pressure ratio CPR is greater than CPR. LIMIT At the same time, reduce the fuel supply to the main combustion chamber to ensure that the compressor pressure ratio does not exceed CPR. LIMIT The fan pressure ratio FPR is greater than FPR. LIMIT At the same time, increase the throat area of ​​the exhaust nozzle to ensure that the fan pressure ratio does not exceed FPR. LIMIT .

[0099] e) The inclusion relationship curves in items a), b), c), and d) need to be corrected and improved as needed based on the accumulated distortion tests of ground stations and high-altitude stations.

[0100] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for controlling the stability margin of a compression component under conditions of overall machine pressure distortion, characterized in that, include: Step S1: Calculate the flow rate W based on the compressor inlet. A25R And the compressor CPR limit line, calculate the compressor pressure ratio limit value CPR under the current operating condition. MAX ; Based on the engine inlet flow rate W A2R And the fan FPR limit line, calculate the fan pressure ratio limit value FPR in the current operating state. MAX ; According to the total inlet pressure P of the engine t2 and the aircraft's total intake pressure P t0 The total pressure loss of the aircraft's air intake, SIGINLET, was calculated. Obtain the converted flow rate W of the engine inlet A2R The relationship between total pressure loss in the intake manifold (SIGINLET) and the inlet distortion index (W); Based on the current state of the aircraft's total pressure loss in the air intake (SIGINLET) and the engine inlet flow rate (W), A2R Based on the aforementioned relationship, the predicted value W of the import distortion index is calculated; then, according to the compressor pressure ratio and fan pressure ratio correction curves under different import distortion index W conditions, the fan pressure ratio correction ΔFPR is calculated respectively. MAX Compressor pressure ratio correction ΔCPR MAX ; Step S2: Correct the compressor pressure ratio ΔCPR MAX The compressor pressure ratio limit (CPR) for the current operating condition. MAX The final compressor pressure ratio limit (CPR) for the current operating state is obtained by making corrections. LIMIT ; through fan pressure ratio correction △FPR MAX The current operating condition fan pressure ratio limit value FPR MAX The final value of the fan pressure ratio limit FPR for the current operating state is obtained by making corrections. LIMIT ; Step S3: When the engine is running: When the compressor pressure ratio CPR is greater than CPR LIMIT At the same time, reduce the fuel supply to the main combustion chamber to ensure that the compressor pressure ratio does not exceed CPR. LIMIT ; When the fan pressure ratio FPR is greater than FPR LIMIT At the same time, increase the throat area of ​​the exhaust nozzle to ensure that the fan pressure ratio does not exceed FPR. LIMIT .

2. The method for controlling the stability margin of the compression component under the condition of overall machine pressure distortion as described in claim 1, characterized in that, In step S1, the converted flow rate W at the engine inlet is calculated. A2R The methods for obtaining it include: Obtain the relative conversion speed of the fan And based on the relative conversion speed of the fan Flow rate W converted from fan inlet A2RJC The calculated flow rate W at the fan inlet is obtained from the relationship calculation. A2RJC ; Based on the influence of fan rotor clearance, adjustable guide vane angle, and Reynolds number, the basic value W of the converted flow rate at the fan inlet is calculated. A2RJC The converted flow rate W at the engine inlet is obtained after correction. A2R .

3. The method for controlling the stability margin of the compression component under the condition of overall machine pressure distortion as described in claim 1, characterized in that, The method for obtaining the fan pressure ratio (FPR) in step S3 includes: Obtain low-pressure converted speed Compared with the fan's external bypass pressure ratio P t13 / P t2 ; Obtain low-pressure converted speed Fan external bypass ratio P t13 / P t2 The relationship curve between the fan pressure ratio (FPR) and the fan pressure ratio (FPR); The fan pressure ratio FPR is calculated.

4. The method for controlling the stability margin of the compression component under the condition of overall machine pressure distortion as described in claim 3, characterized in that, The method for obtaining the compressor pressure ratio (CPR) in step S3 includes: Obtain low-pressure converted speed Fan pressure ratio (FPR) and fan temperature ratio (T) t21 / T t2 The relationship curve, linear interpolation yields the fan temperature ratio T. t21 / T t2 ; According to the fan temperature ratio T t21 / T t2 Fan inlet temperature measurement value T t2 The base value of the total temperature at the fan outlet, T, was calculated. t21JC ; Based on the influence of fan rotor clearance and Reynolds number on fan component efficiency, the baseline value T of the total fan outlet temperature is... t21JC Correction yields the total fan outlet temperature T t21 The corrected formula is: T t21 =T t21JC +DETATCL+DETATREU; Where DETATCL is the fan rotor clearance correction value, and DETATREU is the Reynolds number correction value; Based on the radial distribution difference between the total temperature at the fan outlet and the total temperature at the compressor inlet, the total temperature at the fan outlet T... t21 Correction yields compressor inlet temperature T t25 ; Based on the relative physical rotation speed of high pressure The compressor inlet temperature T t25 The compressor's converted speed was calculated. Obtain the pressure ratio P of the high-pressure compression component t3 / P t21 Obtain the compressor's converted speed High pressure compression component pressure ratio P t3 / P t21 Calculated flow rate W with the internal structure of the intermediate casing A23R Relationship curve; The calculated inlet flow rate W of the intermediate casing is obtained. A23R ; According to formula W A23 =W A23R ×(288.15 / T t21 ) 0.5 ×P t21 The inlet airflow W of the intermediate casing is calculated using / 101.

325. A23 Among them, P t21 This refers to the fan outlet pressure. According to the inlet airflow W inside the intermediate casing A23 The intermediate casing intrinsic loss σ is obtained by calculating the relationship curve between the intermediate casing intrinsic loss and the intermediate casing intrinsic loss. NH , According to formula P t25 =P t21 ×σ NH The compressor inlet pressure P was calculated. t25 ; Based on the compressor outlet pressure measurement value P t3 The compressor inlet pressure P t25 The compressor pressure ratio CPR is calculated using the formula CPR = P t3 / P t25 .

5. The method for controlling the stability margin of the compression component under the condition of overall machine pressure distortion as described in claim 4, characterized in that, The compressor inlet converted flow rate W A25R The calculation formula is: W A25R =W A23 ×(T t25 / 288.15) 0.5 ×101.325 / P t25 。 6. The method for controlling the stability margin of the compression component under the condition of overall machine pressure distortion as described in claim 1, characterized in that, The calculation method for the total pressure loss SIGINLET of the aircraft inlet in step S1 includes: Based on W A2R ~P t2 The / P2 relationship curve, based on the inlet converted flow rate W. A2R The inlet total static pressure ratio P was calculated. t2 / P2 value; Based on the measured static pressure value P2 at the engine inlet and the total static pressure ratio P t2 / P2 value, the total inlet pressure P is calculated. t2 ; The total pressure P ahead is calculated based on the ambient pressure P0 and Mach number signal M provided by the aircraft communication. t0 The calculation formula is: P t0 =P0×(1+(k-1)×M 2 / 2) k / (k-1) ; According to the total inlet pressure P of the engine t2 The total air intake pressure P of the aircraft t0 The total pressure loss SIGINLET of the aircraft air intake is calculated using the formula: SIGINLET = P t2 / P t0 .

7. The method for controlling the stability margin of the compression component under the condition of overall machine pressure distortion as described in claim 1, characterized in that, Current operating condition compressor pressure ratio limit final value CPR LIMIT The calculation formula is: CPR LIMIT =CPR MAX -△CPR MAX ; Current operating state fan pressure ratio limit final value FPR LIMIT The calculation formula is: FPR LIMIT =FPR MAX -△FPR MAX 。

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