A method for evaluating the thrust of an aircraft engine equipped with a convergent nozzle

Abstract

This invention discloses a thrust evaluation method for an aero-engine equipped with a convergent nozzle, comprising the following steps: Step S1: Obtain the compressor characteristic diagram and flight parameter data of the engine; Step S2: Obtain the converted flow rate of the compressor based on the compressor characteristic diagram; Step S3: Determine the actual airflow rate of the compressor based on the converted flow rate, the total temperature at the engine inlet, and the total pressure at the engine inlet; Step S4: Calculate the gas velocity at the outlet of the engine nozzle; Step S5: Calculate the static pressure P9 at the outlet of the nozzle; Step S6: Solve for the engine thrust F based on the actual airflow rate of the compressor, the gas velocity at the outlet, and the static pressure P9 at the outlet of the nozzle. This invention evaluates engine thrust based on the compressor characteristic diagram and flight parameter data, providing a reference for understanding the actual performance of the engine off-site and solving the problem of insufficient thrust measurement methods at current airports.

Description

Technical Field

[0001] This invention relates to the field of engine thrust technology, and in particular to a thrust evaluation method for an aero-engine equipped with a convergent nozzle. Background Technology

[0002] Thrust measurement of aero-engines is a major challenge in current aircraft support operations. Currently, domestic thrust testing primarily involves measurements taken at the factory before the aero-engine is installed, resulting in a significant difference between the measured thrust and the installed thrust. Furthermore, test stands are large, complex, and require numerous supporting systems, leading to high costs and technical demands, making it difficult to configure them individually at airports. To ensure safe and efficient aircraft flight and to evaluate engine performance, relatively accurate thrust measurement data is essential. Therefore, a simple, effective, and relatively high-precision engine thrust measurement method needs to be designed. Summary of the Invention

[0003] The purpose of this invention is to overcome the shortcomings of the prior art and provide a thrust evaluation method for an aero-engine equipped with a convergent nozzle. The method evaluates the engine thrust based on the compressor characteristic diagram and flight parameter data, providing a reference for understanding the actual performance of the engine off-site and solving the problem of insufficient thrust measurement methods at current airports.

[0004] To achieve the above objectives, the technical solution adopted by the present invention is: a thrust evaluation method for an aero-engine equipped with a convergent nozzle, comprising the following steps: Step S1: acquiring the compressor characteristic diagram and flight parameter data of the engine; Step S2: acquiring the converted flow rate of the compressor based on the compressor characteristic diagram. Step S3: Based on the converted flow rate Engine inlet total temperature and engine inlet total pressure Determine the actual air flow rate of the compressor Step S4: Calculate the gas velocity at the engine exhaust nozzle outlet. Step S5: Calculate the static pressure P9 at the nozzle outlet; Step S6: Calculate the static pressure P9 based on the actual airflow of the compressor. Gas velocity at the tailpipe outlet Given the static pressure P9 at the tail nozzle exit, calculate the engine thrust F.

[0005] Preferably, the step of obtaining the converted flow rate of the compressor is... The steps include: Step S201: Calculate the engine inlet total temperature based on the local static temperature. Step S202: Based on rotor speed and engine inlet total temperature Solve for the relative conversion speed Step S203: Based on the compressor characteristic diagram, combined with the relative conversion speed The compressor operating point is obtained from the pressure ratio EPR. Step S204: Based on the compressor operating point The converted flow rate of the compressor can be read from the compressor characteristic diagram. .

[0006] Preferably, the compressor operating point The methods for obtaining this include: interpolating the relative converted speed based on existing data points on the compressor characteristic diagram. Draw the isochronous rotational speed line; based on the pressure ratio EPR, draw a straight line l1 such that l1:Y=EPR; the intersection of l1 and the isochronous rotational speed line is the compressor operating point. .

[0007] Preferably, the actual air flow rate of the compressor It can be expressed as follows:

[0008]

[0009] In the formula: This is the converted flow rate for the low-pressure compressor; Total pressure at the engine inlet; Total engine inlet temperature; Standard atmospheric temperature; Standard atmospheric pressure.

[0010] Preferably, the gas velocity at the engine exhaust nozzle outlet is calculated. This includes: ignoring fuel flow and assuming the engine afterburner is not engaged, the exhaust nozzle is in a critical or supercritical state; ignoring heat loss and assuming the total temperature at the exhaust nozzle outlet and inlet is... If they are equal, then the gas velocity at the nozzle exit is... It can be expressed as follows:

[0011]

[0012] In the formula: The total temperature at the nozzle inlet can be measured by an infrared thermometer or a temperature measuring device placed at the nozzle inlet. This is the compressor outlet gas velocity coefficient; R is the specific heat ratio of the fuel gas, taken as 1.33; R is the gas constant.

[0013] Preferably, the gas velocity at the engine exhaust nozzle outlet is calculated. This includes: ignoring fuel flow and assessing the critical or supercritical state of the exhaust nozzle when engine afterburner is engaged; and ignoring heat loss to obtain the engine's afterburner temperature. The gas velocity at the tailpipe outlet... It can be expressed as follows:

[0014]

[0015] In the formula: The afterburner temperature of the engine; k is the exhaust gas velocity coefficient at the nozzle exit. Let R be the specific heat ratio of the fuel gas after the booster is turned on, taken as 1.27; R is the gas constant.

[0016] Preferably, calculating the static pressure P9 at the nozzle outlet includes: the nozzle operating in a critical or supercritical state; assuming the total pressure at the mixer inner inlet and the total pressure at the mixer outer bypass inlet are the same; ignoring the total pressure loss of the outer bypass flow, the total pressure at the mixer inlet is equal to the total pressure at the fan outlet, i.e. The static pressure P9 at the nozzle exit is expressed by the following formula:

[0017]

[0018] In the formula: Total pressure at the mixer inlet; This represents the total pressure loss from the mixer inlet to the tail nozzle outlet. Specific heat ratio of fuel gas; This is the Mach number at the nozzle exit.

[0019] Preferably, the calculation of the engine thrust F includes: neglecting the engine's fuel flow rate and bleed air flow rate; assuming that the gas flow rate at the exhaust nozzle outlet and the air flow rate through the fan are equal; then the engine thrust F is expressed by the following formula:

[0020]

[0021] In the formula: This represents the actual airflow rate of the compressor. V is the exhaust gas velocity at the nozzle exit; P9 is the flight speed; P0 is the static pressure at the nozzle exit; and A9 is the nozzle exit area.

[0022] Preferably, the flight parameter data includes: fan outlet pressure. Total pressure at the fan inlet Tail nozzle inlet total temperature Flight speed V, flight Mach number Ma, low-pressure rotor speed .

[0023] Compared with the prior art, the present invention has the following advantages:

[0024] 1. This invention evaluates engine thrust based on compressor characteristic diagrams and flight parameter data, providing a reference for understanding the actual performance of engines off-site and solving the problem of insufficient thrust measurement methods at current airports.

[0025] 2. This invention uses the low-pressure compressor characteristic diagram, existing flight parameter data, and the total temperature of the exhaust gas at the compressor tail nozzle outlet to perform inference calculations. The calculation method is simple and rapid, which improves the efficiency of thrust measurement.

[0026] The present invention will now be described in further detail with reference to the accompanying drawings and embodiments. Attached Figure Description

[0027] Figure 1 This is a schematic diagram illustrating how the present invention locates the operating point on a low-pressure compressor characteristic diagram. Detailed Implementation

[0028] like Figure 1 As shown, this invention discloses a thrust evaluation method for an aero-engine equipped with a convergent nozzle, comprising the following steps:

[0029] Step S1: Obtain the compressor characteristic diagram and flight parameter data of the engine;

[0030] The flight parameter data includes: fan outlet pressure. Total pressure at the fan inlet Tail nozzle inlet total temperature Flight speed V, flight Mach number Ma, low-pressure rotor speed .

[0031] Step S2: Obtain the converted flow rate of the compressor based on the compressor characteristic diagram. ;

[0032] The converted flow rate of the compressor is obtained. The steps include:

[0033] Step S201: Calculate the engine inlet total temperature based on the local static temperature. ;

[0034] Local calm temperature and static pressure The calculation method is as follows:

[0035] When H ≤ 11km:

[0036] (1)

[0037] (2)

[0038] When altitude H > 11km:

[0039] (3)

[0040] (4)

[0041] Engine inlet total temperature It can be expressed as follows:

[0042] (5)

[0043] In the formula: The local static temperature; k air is the specific heat ratio of air; Ma is the flight Mach number.

[0044] Engine inlet (fan inlet) total pressure It can be expressed as follows:

[0045] (6)

[0046] In the formula: The local static pressure.

[0047] Step S202: Based on rotor speed and engine inlet total temperature Solve for the relative conversion speed ;

[0048] The relative conversion speed It can be expressed as follows:

[0049] (7)

[0050] In the formula: This refers to the rotor speed; Total engine inlet temperature; The standard atmospheric temperature is (generally taken as 288.15K).

[0051] Step S203: Based on the compressor characteristic diagram, combined with the relative conversion speed The compressor operating point is obtained from the pressure ratio EPR. ;

[0052] The compressor operating point The methods for obtaining it include:

[0053] Based on the interpolation of existing data points on the compressor characteristic diagram, the relative converted speed is: isochronous rotation lines;

[0054] Draw a straight line l1 based on the pressure ratio EPR, such that l1:Y=EPR;

[0055] The point where l1 intersects with the isochronous rotation line is the compressor operating point. .

[0056] Step S204: Based on the compressor operating point The converted flow rate of the compressor can be read from the compressor characteristic diagram. .

[0057] Step S3: Based on the converted flow rate Engine inlet total temperature and engine inlet total pressure Determine the actual air flow rate of the compressor ;

[0058] The actual air flow rate of the compressor It can be expressed as follows:

[0059] (8)

[0060] In the formula: This is the converted flow rate for the low-pressure compressor; Total air pressure at the engine inlet; Total air temperature at the engine inlet; Standard atmospheric temperature (288.15K); Standard atmospheric pressure (101325 Pa).

[0061] Step S4: Calculate the gas velocity at the engine exhaust nozzle outlet. ;

[0062] Calculate the gas velocity at the engine exhaust nozzle outlet when the engine afterburner is not engaged. include:

[0063] Ignoring fuel flow, the tail nozzle is in a critical or supercritical state;

[0064] Neglecting heat loss, assume the total temperature at the nozzle outlet and the total temperature at the nozzle inlet are... equal;

[0065] The gas velocity at the engine exhaust nozzle outlet It can be expressed as follows:

[0066] (9)

[0067] In the formula: The total temperature at the nozzle inlet can be measured by an infrared thermometer or a temperature measuring device placed at the nozzle inlet. This is the exhaust gas velocity coefficient at the tailpipe outlet; R is the specific heat ratio of the fuel gas, taken as 1.33; R is the gas constant.

[0068] When the engine afterburner is activated, calculate the gas velocity at the engine exhaust nozzle outlet. include:

[0069] Ignoring fuel flow, the tail nozzle is in a critical or supercritical state;

[0070] Ignoring heat loss, obtain the engine's afterburner temperature. ;

[0071] The gas velocity at the tailpipe outlet It can be expressed as follows:

[0072] (10)

[0073] In the formula: The afterburner temperature of the engine can be measured by an infrared thermometer or a temperature measuring device placed at the tail nozzle outlet, or it can be estimated based on the engine manual. This is the exhaust gas velocity coefficient at the tailpipe outlet; Let R be the specific heat ratio of the fuel gas after the booster is turned on, taken as 1.27; R is the gas constant.

[0074] Step S5: Calculate the static pressure P9 at the tail nozzle exit;

[0075] The calculation of the static pressure P9 at the tailpipe outlet includes:

[0076] The tail nozzle operates in a critical or supercritical state.

[0077] Assume the total pressure at the inlet of the mixer's inner chamber is the same as the total pressure at the inlet of the mixer's outer chamber;

[0078] Ignoring the total pressure loss of the bypass duct flow, the total pressure at the mixer inlet is made equal to the total pressure at the fan outlet, i.e. ;

[0079] The static pressure P9 at the tailpipe outlet is related to the total pressure. The relationship is shown in equation (11):

[0080] (11)

[0081] The static pressure P9 at the nozzle exit can be calculated using the following formula:

[0082] (12)

[0083] In the formula: Total pressure at the mixer inlet; k is the total pressure loss from the mixer inlet to the tailpipe outlet. gas Specific heat ratio of fuel gas; This is the Mach number at the nozzle exit. Since the nozzle operates in a critical or supercritical state, therefore... It equals 1.

[0084] Step S6: Based on the actual air flow rate of the compressor Gas velocity at the tailpipe outlet Given the static pressure P9 at the tail nozzle exit, calculate the engine thrust F.

[0085] The calculation of the engine thrust F includes:

[0086] Ignore engine fuel flow and bleed air flow;

[0087] Assume that the gas flow rate at the tailpipe outlet is equal to the air flow rate through the fan;

[0088] The engine thrust F is then expressed by the following formula:

[0089] (13)

[0090] In the formula: This represents the actual airflow rate of the compressor. V is the exhaust gas velocity at the nozzle exit; P9 is the flight speed; P0 is the static pressure at the nozzle exit; A9 is the nozzle exit area, which is obtained from the engine operating manual.

[0091] Using a numerical model of a hybrid turbofan engine as the verification object, the thrust calculated by the method of this application is compared with the simulation results of the numerical model.

[0092] Comparative Example 1

[0093] Given the engine operating conditions as follows: flight Mach number Ma = 0.1, flight altitude 0 km, ambient temperature under standard atmospheric conditions (static temperature 288.15 K, static pressure 101325 Pa), and the relative physical speed of the high-pressure rotor... The relative physical speed of the low-pressure rotor is 0.985. The total pressure loss of this engine from the mixer to the exhaust nozzle is 0.9146. The value is 0.93; firstly, the thrust is calculated to be 10310.8 DaN using the simulation program provided by the manufacturer, and then the fan outlet pressure is read from the flight parameter data generated by the simulation program. The total temperature at the nozzle inlet is 206083.4 Pa. It is 519.97K.

[0094] Example 1

[0095] Environmental parameters are calculated based on atmospheric environment, including static temperature. 288.15K, static pressure The total pressure at the engine inlet is 101325 Pa. Calculate the total pressure at the engine inlet based on the relationship between the total pressure and static pressure at the engine inlet. air specific heat ratio k air Taking 1.4, and the flight Mach number Ma is 0.1, then according to equation (6), we can obtain:

[0096]

[0097] Calculate the engine inlet total temperature based on the relationship between engine inlet total temperature and static temperature. air specific heat ratio k air Taking 1.4, and the flight Mach number Ma as 0.1, then according to equation (5), we can obtain:

[0098]

[0099] Based on the relative physical speed of the low-pressure rotor and engine inlet total temperature Solve for the relative conversion speed ;

[0100] According to equation (7), we can obtain:

[0101]

[0102] That is, relative conversion speed It is 0.9137;

[0103] Based on the compressor characteristic diagram, combined with the relative conversion speed The compressor operating point is obtained by interpolation based on the fan pressure ratio (EPR). ;

[0104] The pressure ratio EPR is based on the fan outlet pressure. and engine inlet total pressure Compared to what we get, that is:

[0105]

[0106] Obtain the fan outlet pressure based on flight data. And the total engine inlet pressure obtained by equation (6) Then the pressure ratio EPR is 2.0197;

[0107] like Figure 1 As shown, based on the compressor operating point The converted flow rate of the compressor can be read from the compressor characteristic diagram. The flow rate is 275.11 kg / s, based on the converted flow rate. Engine inlet total temperature and engine inlet total pressure Determine the actual air flow rate of the compressor ;

[0108] According to equation (8), we can obtain:

[0109]

[0110] That is, the actual air flow rate of the compressor. It is 275.39 kg / s;

[0111] Calculate the gas velocity at the compressor outlet when the engine afterburner is not engaged. Ignoring fuel flow, the exhaust nozzle is in a critical or supercritical state. Since the exhaust nozzle is convergent, the exit velocity coefficient is... Assuming the specific heat ratio of the fuel gas is 1, the specific heat ratio of the fuel gas is assumed to be 1. The value is 1.33, the gas constant R is 287.06, and heat loss from the nozzle, afterburner, etc. is ignored. Assume the total temperature at the nozzle outlet and the total temperature at the nozzle inlet are... Equal to each other, the gas velocity at the engine tailpipe outlet can be obtained according to equation (9). for:

[0112]

[0113] The gas velocity at the tailpipe outlet It is 412.8 m / s;

[0114] At this point, the exhaust nozzle is operating in a critical or supercritical state. According to the design principles of mixed-emission turbofan engines, the ratio of the total pressure at the inlet of the mixer inner chamber to the total pressure at the inlet of the mixer outer chamber is between 0.98 and 1.02. Assuming the ratio of the total pressure at the inlet of the mixer inner chamber to the total pressure at the inlet of the mixer outer chamber is 1, then the total pressure at the mixer inlet is the same as the total pressure at the fan outlet. Determine the total pressure loss from the mixer to the exhaust nozzle outlet based on the engine operating manual or experience. ,Pick Since it is 0.93, according to equation (12), the static pressure P9 at the nozzle outlet is:

[0115]

[0116] Ignoring fuel flow and bleed air flow, assuming the gas flow at the exhaust nozzle outlet is equal to the air flow through the fan, and based on the actual air flow of the compressor... Gas velocity at the tailpipe outlet Given the static pressure P9 at the tail nozzle exit, calculate the engine thrust F based on the flight Mach number Ma and static temperature. The flight speed V can be obtained, and the tail nozzle exit area A9 is obtained from the engine instruction manual. Then, according to equation (13), the engine thrust F can be obtained as follows:

[0117]

[0118] The engine thrust F calculated using the thrust evaluation method of this application is 10356 DaN, while the thrust F calculated through numerical model simulation is 10310.8 DaN. Therefore, the engine thrust error calculated by this application is -0.44%, which is within the preset error range and meets the requirements for engineering use.

[0119] Comparative Example 2

[0120] Given the engine operating conditions as follows: flight Mach number Ma = 0.7, flight altitude 6 km, and high-pressure rotor relative physical speed... The relative physical speed of the low-pressure rotor is 0.91. The thrust F was first calculated to be 3310.5 DaN using a thrust calculation program, and then the fan outlet pressure was read from the flight parameter data generated by the simulation program. The pressure is 115953.5 Pa, and the nozzle inlet temperature is... The total pressure loss of this engine from the mixer to the exhaust nozzle is 451.04K. It remains unchanged at 0.93.

[0121] Example 2

[0122] Environmental parameters are calculated based on atmospheric environment, including static temperature. 249.15K, static pressure The total pressure at the engine inlet is 47165 Pa. Calculate the total pressure at the engine inlet based on the relationship between the total pressure and static pressure at the engine inlet. air specific heat ratio k air Taking 1.4, and the flight Mach number Ma is 0.7, then according to equation (6), we can obtain:

[0123]

[0124] That is, the total pressure at the engine inlet The Pa is 65456.7.

[0125] Calculate the engine inlet total temperature based on the relationship between engine inlet total temperature and static temperature. air specific heat ratio k air Taking 1.4, and the flight Mach number Ma is 0.7, then according to equation (5), we can obtain:

[0126]

[0127] That is, the total temperature at the engine inlet It is 273.73K;

[0128] Based on rotor speed and engine inlet total temperature Solve for the relative conversion speed ;

[0129] According to equation (7), we can obtain:

[0130]

[0131] That is, relative conversion speed It is 0.8266;

[0132] Based on the compressor characteristic diagram and the converted speed The compressor operating point is obtained by interpolation based on the pressure ratio EPR. ;

[0133] The pressure ratio EPR is based on the fan outlet pressure. and total imported pressure Compared to what we get, that is:

[0134]

[0135] Obtain the fan outlet pressure based on flight data. And the total engine inlet pressure obtained by equation (6) Then the pressure ratio EPR is 1.771;

[0136] like Figure 1 As shown, based on the compressor operating point The converted flow rate of the compressor can be read from the compressor characteristic diagram. The flow rate is 259.3 kg / s, based on the converted flow rate. Engine inlet total temperature and engine inlet total pressure Determine the actual air flow rate of the compressor ;

[0137] According to equation (8), we can obtain:

[0138]

[0139] That is, the actual air flow rate of the compressor. It is 163.3 kg / s;

[0140] Calculate the gas velocity at the compressor outlet when the engine afterburner is not engaged. Ignoring fuel flow, the exhaust nozzle is in a critical or supercritical state. Since the exhaust nozzle is convergent, the exit velocity coefficient is... Assuming the specific heat ratio of the fuel gas is 1, the specific heat ratio of the fuel gas is assumed to be 1. The gas constant R is 287.06, and heat loss from the exhaust nozzle and afterburner is ignored. Assuming the total exhaust temperature at the exhaust nozzle outlet and the total exhaust gas temperature at the exhaust nozzle inlet are... Equal to each other, the gas velocity at the engine tail nozzle outlet can be obtained according to equation (9). for:

[0141]

[0142] The gas velocity at the tailpipe outlet It is 384.46 m / s;

[0143] At this point, the exhaust nozzle is operating in a critical or supercritical state. According to the design principles of mixed-emission turbofan engines, the ratio of the total pressure at the inlet of the mixer inner chamber to the total pressure at the inlet of the mixer outer chamber is between 0.98 and 1.02. Assuming the ratio of the total pressure at the inlet of the mixer inner chamber to the total pressure at the inlet of the mixer outer chamber is 1, then the total pressure at the mixer inlet is the same as the total pressure at the fan outlet. Determine the total pressure loss from the mixer to the exhaust nozzle outlet based on the engine operating manual or experience. ,Pick Since it is 0.93, according to equation (12), the static pressure P9 at the nozzle outlet is:

[0144]

[0145] Ignoring fuel flow and bleed air flow, assuming the gas flow at the exhaust nozzle outlet is equal to the air flow through the fan, and based on the actual air flow of the compressor... Gas velocity at the tailpipe outlet Given the static pressure P9 at the tail nozzle exit, calculate the engine thrust F based on the flight Mach number Ma and static temperature. The flight speed V can be obtained, and the tail nozzle exit area A9 is obtained from the engine instruction manual. Then, according to equation (13), the engine thrust F can be obtained as follows:

[0146]

[0147] The engine thrust F calculated using the thrust evaluation method of this application is 3453.3 DaN, while the thrust F calculated through numerical model simulation is 3310.5 DaN. Therefore, the engine thrust error calculated by this application is 4.31%, which is within the preset error range and meets the requirements for engineering use.

[0148] The same method was used to calculate working condition 3-6, and the results are shown in Table 1:

[0149] Table 1 Comparison of Thrust Calculation Results

[0150] Serial Number Ma Alt High-voltage rotor relative speed Low-pressure rotor relative speed Simulation calculation of thrust This method calculates thrust. error 1 0.1 0 0.9850 0.9146 10310.8 10356 0.44% 2 0.7 6 0.9100 0.8057 3310.5 3453.3 4.31% 3 0.5 4 0.9200 0.8224 4298 4435.6 3.20% 4 0.7 8 0.9200 0.8490 3290 3344.3 1.65% 5 0.6 3 0.9850 0.8510 4809.1 4928.4 2.48% 6 0.8 3 0.9600 0.8512 5368.3 5575.5 3.86%

[0151] As shown in Table 1, under multiple operating conditions, the maximum error of the engine thrust F calculated using the method of this application for a convergent nozzle aero-engine with a service ceiling of 9 km is 4.31%, which is within the preset error range and meets the requirements for engineering use. Moreover, the calculation method of this application is simple and fast, which improves the efficiency of thrust calculation.

[0152] The above description is merely a preferred embodiment of the present invention and does not constitute any limitation on the present invention. Any simple modifications, alterations, or equivalent structural transformations made to the above embodiments based on the technical essence of the present invention shall still fall within the protection scope of the present invention.

Claims

1. A thrust evaluation method for an aero-engine equipped with a convergent nozzle, characterized in that: Includes the following steps: Step S1: Obtain the compressor characteristic diagram and flight parameter data of the engine; Step S2: Obtain the converted flow rate of the compressor based on the compressor characteristic diagram. ; Step S3: Based on the converted flow rate Engine inlet total temperature and engine inlet total pressure Determine the actual air flow rate of the compressor ; Step S4: Calculate the gas velocity at the engine exhaust nozzle outlet. ; Step S5: Calculate the static pressure P9 at the tail nozzle exit; Step S6: Based on the actual air flow rate of the compressor Gas velocity at the tailpipe outlet Given the static pressure P9 at the tail nozzle exit, calculate the engine thrust F.

2. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 1, characterized in that: The converted flow rate of the compressor is obtained. The steps include: Step S201: Calculate the engine inlet total temperature based on the local static temperature. ; Step S202: Based on rotor speed and engine inlet total temperature Solve for the relative conversion speed ; Step S203: Based on the compressor characteristic diagram, combined with the relative conversion speed The compressor operating point is obtained from the pressure ratio EPR. ; Step S204: Based on the compressor operating point The converted flow rate of the compressor can be read from the compressor characteristic diagram. .

3. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 2, characterized in that: The compressor operating point The methods for obtaining it include: Based on the interpolation of existing data points on the compressor characteristic diagram, the relative converted speed is: isochronous rotation lines; Draw a straight line l1 based on the pressure ratio EPR, such that l1:Y=EPR; The point where l1 intersects with the isochronous rotation line is the compressor operating point. .

4. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 1, characterized in that: The actual air flow rate of the compressor It can be expressed as follows: In the formula: This is the converted flow rate for the low-pressure compressor; Total pressure at the engine inlet; Total engine inlet temperature; Standard atmospheric temperature; Standard atmospheric pressure.

5. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 1, characterized in that: Calculate the gas velocity at the engine exhaust nozzle outlet. include: Ignoring fuel flow, when the engine afterburner is not engaged, the exhaust nozzle is in a critical or supercritical state. Neglecting heat loss, assume the total temperature at the nozzle outlet and the total temperature at the nozzle inlet are... equal; The gas velocity at the engine exhaust nozzle outlet It can be expressed as follows: In the formula: The total temperature at the nozzle inlet can be measured by an infrared thermometer or a temperature measuring device placed at the nozzle inlet. This is the exhaust gas velocity coefficient at the tailpipe outlet; R is the specific heat ratio of the fuel gas, taken as 1.33; R is the gas constant.

6. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 1, characterized in that: Calculate the gas velocity at the engine exhaust nozzle outlet. include: Ignoring fuel flow, when the engine afterburner is engaged, the exhaust nozzle is in a critical or supercritical state. Ignoring heat loss, obtain the engine's afterburner temperature. ; The gas velocity at the tailpipe outlet It can be expressed as follows: In the formula: The afterburner temperature of the engine; This is the exhaust gas velocity coefficient at the tailpipe outlet; Let R be the specific heat ratio of the fuel gas after the booster is turned on, taken as 1.27; R is the gas constant.

7. The thrust evaluation method for an aero-engine equipped with a convergent nozzle according to claim 1, characterized in that: The calculation of the static pressure P9 at the tailpipe outlet includes: The tail nozzle operates in a critical or supercritical state. Assume the total pressure at the inlet of the mixer's inner chamber is the same as the total pressure at the inlet of the mixer's outer chamber; Ignoring the total pressure loss in the bypass duct, the total pressure at the mixer inlet is equal to the total pressure at the fan outlet, i.e. ; The static pressure P9 at the nozzle exit is expressed by the following formula: In the formula: Total pressure at the mixer inlet; This represents the total pressure loss from the mixer inlet to the tail nozzle outlet. Specific heat ratio of fuel gas; This is the Mach number at the nozzle exit.

8. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 1, characterized in that: The calculation of the engine thrust F includes: Ignore engine fuel flow and bleed air flow; Assume that the gas flow rate at the tailpipe outlet is equal to the air flow rate through the fan; The engine thrust F is then expressed by the following formula: In the formula: This represents the actual airflow rate of the compressor. V is the exhaust gas velocity at the nozzle exit; P9 is the flight speed; P0 is the static pressure at the nozzle exit; and A9 is the nozzle exit area.

9. The thrust evaluation method for an aero-engine equipped with a convergent tail nozzle according to claim 1, characterized in that: The flight parameter data includes: fan outlet pressure. Total pressure at the fan inlet Tail nozzle inlet total temperature Flight speed V, flight Mach number Ma, low-pressure rotor speed .

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