CN108182297B - Rocket engine combustion chamber efficiency analysis method and system - Google Patents

Rocket engine combustion chamber efficiency analysis method and system Download PDF

Info

Publication number
CN108182297B
CN108182297B CN201710766220.9A CN201710766220A CN108182297B CN 108182297 B CN108182297 B CN 108182297B CN 201710766220 A CN201710766220 A CN 201710766220A CN 108182297 B CN108182297 B CN 108182297B
Authority
CN
China
Prior art keywords
combustion chamber
engine
throat
calculating
calculation module
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201710766220.9A
Other languages
Chinese (zh)
Other versions
CN108182297A (en
Inventor
杨成虎
林庆国
张志远
刘昌国
刘犇
赵婷
姚锋
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Shanghai Institute of Space Propulsion
Original Assignee
Shanghai Institute of Space Propulsion
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Shanghai Institute of Space Propulsion filed Critical Shanghai Institute of Space Propulsion
Priority to CN201710766220.9A priority Critical patent/CN108182297B/en
Publication of CN108182297A publication Critical patent/CN108182297A/en
Application granted granted Critical
Publication of CN108182297B publication Critical patent/CN108182297B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/08—Thermal analysis or thermal optimisation

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Theoretical Computer Science (AREA)
  • General Physics & Mathematics (AREA)
  • Evolutionary Computation (AREA)
  • Computer Hardware Design (AREA)
  • General Engineering & Computer Science (AREA)
  • Pure & Applied Mathematics (AREA)
  • Mathematical Optimization (AREA)
  • Mathematical Analysis (AREA)
  • Computational Mathematics (AREA)
  • Testing Of Engines (AREA)
  • Combined Controls Of Internal Combustion Engines (AREA)

Abstract

The invention discloses a rocket engine combustion chamber efficiency analysis method and a rocket engine combustion chamber efficiency analysis system. The invention more accurately realizes the analysis of the efficiency of the combustion chamber and provides more accurate basis for the performance evaluation comparison of different working conditions and different engines of the same engine.

Description

Rocket engine combustion chamber efficiency analysis method and system
Technical Field
The invention relates to the field of rocket engines, in particular to a method and a system for analyzing the efficiency of a combustion chamber of a rocket engine.
Background
The combustion chamber efficiency is an important parameter for evaluating the performance of the engine, and in the process of developing the engine, the combustion chamber efficiency and the nozzle efficiency of the engine need to be obtained through engine hot test examination, wherein the nozzle efficiency is obtained through calculation based on the specific impulse of the engine and the combustion chamber efficiency, so that the combustion chamber efficiency of the engine is accurately obtained, and the method is very important for accurately evaluating the performance of the engine and improving or determining the design of the engine.
At present, in the hot test run of an engine, the efficiency of a combustion chamber cannot be directly measured, but is obtained by measuring parameters such as the diameter of the throat part of the engine before the test run, the pressure of the combustion chamber, the flow rate of a propellant and the like in the test run process and then calculating according to the following formula:
Figure BDA0001394160730000011
from the calculation formula, the combustion chamber efficiency has no relation with the engine operating temperature. However, due to the effect of expansion with heat and contraction with cold of the material, when the actual engine works, the temperature of the throat part reaches over 1000 ℃, the diameter of the throat part is generally larger than the diameter value measured at normal temperature before test run, so that the flow rate of the propellant generating the same combustion chamber pressure is larger than that before expansion, therefore, the efficiency of the combustion chamber obtained by the method is generally lower than the true value, and if the temperature of the throat part reaches 1200 ℃, the linear expansion coefficient of the material is 8 e-6/DEG C, the result is calculated by the traditional method to deviate from the true value by nearly 2%. On the other hand, the throat temperature stabilization time is generally longer than the flow and pressure stabilization time when the engine works, the diameter of the throat of the combustion chamber is changed before the throat temperature is stabilized, and the combustion chamber efficiency obtained at the stage is not a true value. It is noted that the higher the throat temperature, the greater the deviation of the combustion chamber efficiency calculated by the conventional method from the true value, which in turn causes the deviation of the nozzle efficiency calculated based on the combustion chamber efficiency from the true value in the same proportion. Therefore, when the working temperature of the same engine is different due to the difference of working conditions or different working temperatures of different engines are different, the combustion chamber efficiency and the nozzle efficiency obtained by the traditional method cannot truly reflect the difference, and therefore the improvement and confirmation of the design state of the engine are misled.
Disclosure of Invention
In order to overcome the defects in the prior art, the invention provides a method and a system for analyzing the efficiency of a combustion chamber of a rocket engine, which provide accurate basis for evaluating and comparing the performance of the same engine under different working conditions and different engines.
The purpose of the invention is realized by the following technical scheme: a rocket engine combustion chamber efficiency analysis system comprises
Data acquisition module for performing normal temperature T0Lower engine throat diameter Dt0Corresponding linear expansion coefficient alpha of combustion chamber material and engine combustion chamber pressure P during engine thermal testc(t), oxidizer and fuel mass flow rate qmo(t) and qmf(T) and throat temperature Tt(t) collecting;
the data average value calculation module is used for obtaining the average value of each parameter of the engine through a preset algorithm to obtain the average value of the pressure in the combustion chamber
Figure BDA0001394160730000021
Average mass flow of oxidant and fuel
Figure BDA0001394160730000022
And
Figure BDA0001394160730000023
mean value of throat temperature
Figure BDA0001394160730000024
A throat diameter calculation module for, according to a preset formula:
Figure BDA0001394160730000025
diameter D of throat portion of engine during operationtCalculating (1); wherein alpha is the linear expansion coefficient of the combustion chamber base material at the corresponding temperature;
a mixing ratio calculation module for, according to a preset formula: mixing ratio
Figure BDA0001394160730000026
Mixing ratio r of oxidant and fuel flowmCalculating (1);
an actual characteristic velocity calculation module for calculating a velocity based on the predictionThe formula is as follows:
Figure BDA0001394160730000027
proceed to the actual characteristic speed C*Calculating (1); wherein D istIs the temperature corrected throat diameter;
theoretical characteristic velocity calculation module using combustion chamber pressure
Figure BDA00013941607300000210
And mixing ratio rmCalculating theoretical characteristic speed according to the principle of minimum Gibbs free energy
Figure BDA0001394160730000028
An engine combustion chamber efficiency calculation module for passing a predetermined formula
Figure BDA0001394160730000029
Efficiency η of combustion chamber of enginecAnd (4) calculating.
The invention also provides a rocket engine combustion chamber efficiency analysis method, which comprises the following steps:
step 1, collecting normal temperature T through a data collection module0Lower engine throat diameter Dt0Inquiring the linear expansion coefficient alpha of the corresponding combustion chamber material;
step 2, carrying out hot test of the engine, and acquiring the pressure P of the combustion chamber of the engine through a data acquisition modulec(t), oxidizer and fuel mass flow rates qmo(t) and qmf(T), throat temperature Tt(t)；
Step 3, obtaining an arithmetic mean value of each parameter of the engine through a data mean value calculation module to obtain a mean value of the pressure in the combustion chamber
Figure BDA0001394160730000031
Average mass flow of oxidant and fuel
Figure BDA0001394160730000032
And
Figure BDA0001394160730000033
mean value of throat temperature
Figure BDA0001394160730000034
Step 4, calculating the throat diameter D of the engine during working through a throat diameter calculation module by utilizing the average value of the throat temperaturet：
Figure BDA0001394160730000035
Wherein alpha is the linear expansion coefficient of the combustion chamber base material at the corresponding temperature;
and calculating the mixing ratio r by a mixing ratio calculation module according to the average flow of the oxidant and the fuelm：
Figure BDA0001394160730000036
Step 5, calculating the actual characteristic speed C through an actual characteristic speed calculation module*：
Figure BDA0001394160730000037
Wherein D istIs the temperature corrected throat diameter;
step 6, utilizing the pressure of the combustion chamber
Figure BDA0001394160730000038
And mixing ratio rmCalculating the theoretical characteristic speed according to the minimum Gibbs free energy principle by a theoretical characteristic speed calculation module
Figure BDA0001394160730000039
Step 7, calculating the efficiency eta of the engine combustion chamber through an engine combustion chamber efficiency calculation modulec：
Figure BDA00013941607300000310
Wherein the temperature-corrected throat diameter is calculated by adopting a linear expansion coefficient
Figure BDA00013941607300000311
When averaging all the parameters, in order to ensure that the parameters are taken from the same engine state, the average value calculation interval is the same time interval, and when averaging the parameters in the engine test run process, the length of the test run program is required to ensure that the parameters such as the pressure of a combustion chamber, the flow of oxidant and fuel, the throat temperature and the like have enough stable sections, and the length of the stable sections is not less than 5 s.
Compared with the prior art, the invention has the following beneficial effects:
in engine design, in order to obtain higher specific impact performance, it is necessary to organize more sufficient combustion to generate higher temperatures, and on the other hand, in order to reduce the requirements for high temperature resistant materials, it is desirable that the lower the engine operating temperature, the better. Therefore, the working temperature of the engine needs to be reduced as much as possible on the premise of not influencing the performance of the engine, and the performance of the engine needs to be evaluated through the efficiency of the combustion chamber during test of the short nozzle. According to the analysis, if the influence of the temperature on the throat diameter of the combustion chamber is not considered, the performance of the engine at higher temperature can be underestimated, so that misjudgment is formed on the temperature reduction design of the engine, and lower specific impulse performance of the engine can be obtained only when a full-size high-altitude simulation thermal test is carried out later. By the aid of the method for analyzing the efficiency of the combustion chamber, the performance of the engine can be accurately evaluated in a short nozzle examination stage by considering the influence of temperature, and development cost and period are remarkably saved.
Drawings
FIG. 1 is a schematic diagram of the change of the diameter of the throat part of the engine in the embodiment of the invention.
FIG. 2 is a graphical representation of engine operating parameters in accordance with an embodiment of the present invention.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that variations and modifications can be made by persons skilled in the art without departing from the spirit of the invention. All falling within the scope of the present invention. In the description of the present invention, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, are not to be construed as limiting the present invention.
The embodiment of the invention provides a rocket engine combustion chamber efficiency analysis system, which comprises
Data acquisition module for performing normal temperature T0Lower engine throat diameter Dt0Corresponding linear expansion coefficient alpha of combustion chamber material and engine combustion chamber pressure P during engine thermal testc(t), oxidizer and fuel mass flow rate qmo(t) and qmf(T) and throat temperature Tt(t) collecting;
the data average value calculation module is used for obtaining the average value of each parameter of the engine through a preset algorithm to obtain the average value of the pressure in the combustion chamber
Figure BDA0001394160730000041
Average mass flow of oxidant and fuel
Figure BDA0001394160730000042
And
Figure BDA0001394160730000043
mean value of throat temperature
Figure BDA0001394160730000044
A throat diameter calculation module for, according to a preset formula:
Figure BDA0001394160730000045
diameter D of throat portion of engine during operationtCalculating (1); wherein alpha is the linear expansion coefficient of the combustion chamber base material at the corresponding temperature;
a mixing ratio calculation module for, according to a preset formula: mixing ratio
Figure BDA0001394160730000046
Mixing ratio r of oxidant and fuel flowmCalculating (1);
an actual characteristic velocity calculation module, configured to:
Figure BDA0001394160730000047
proceed to the actual characteristic speed C*Calculating (1); wherein D istIs the temperature corrected throat diameter;
theoretical characteristic velocity calculation module using combustion chamber pressure
Figure BDA0001394160730000048
And mixing ratio rmCalculating theoretical characteristic speed according to the principle of minimum Gibbs free energy
Figure BDA0001394160730000049
An engine combustion chamber efficiency calculation module for passing a predetermined formula
Figure BDA0001394160730000051
Efficiency η of combustion chamber of enginecAnd (4) calculating. The throat diameter after temperature correction is obtained by adopting linear expansion coefficient calculation
Figure BDA0001394160730000052
When averaging all parameters, in order to ensure that the parameters are taken from the same engine state, the average value calculation interval is the same time interval, and when averaging the parameters in the engine test run process, the test run program length is required to ensure the pressure of a combustion chamber, the oxidant and the fuel flowAnd the throat temperature and other parameters have enough stable sections, and the length of each stable section is not less than 5 s.
FIG. 1 is a schematic diagram of a change in engine combustion chamber profile relative to a normal temperature condition under operating conditions. The solid line part of the graph is the profile of the combustion chamber at normal temperature, and the ambient temperature T can be measured before the test is started0Lower throat diameter Dt0. The dotted line is the state after the combustion chamber expands to generate radial and axial displacement under the working state, and the diameter of the throat part after expansion is D in the figuret。
Fig. 2 is a graph of engine combustion chamber pressure, propellant flow rate and throat temperature measured during the test, for convenience of illustration, throat temperature is plotted against the actual parameter of 0.1 and pressure data is plotted against the actual parameter of 10 times. Taking the average value of each parameter stable section between two virtual lines, and calculating according to the average value of the throat temperature and the linear expansion coefficient alpha of the material to obtain the throat diameter of
Figure BDA0001394160730000053
And calculating the characteristic speed and efficiency of the combustion chamber by using the throat diameter of the working state.
Examples
Taking a certain engine combustion chamber efficiency analysis as an example,
step 1, collecting normal temperature T of an engine through a data collection module0Throat diameter D at (25 ℃ C.)t018.00mm and the material linear expansion coefficient α is found to be 6 e-6/deg.c.
Step 2, acquiring data through a data acquisition module during engine test, wherein a relevant parameter curve is shown in figure 2, each parameter of the engine is stable after 40s according to the test parameter curve (figure 2), and taking the parameter in an interval of 40 s-48 s (between two vertical dotted lines in the figure);
step 3, carrying out average calculation on the obtained parameters through a data average calculation module to obtain the average value of the throat temperature
Figure BDA0001394160730000054
Average oxidant flow
Figure BDA0001394160730000055
Average fuel flow
Figure BDA0001394160730000056
Mean value of combustion chamber pressure
Figure BDA0001394160730000057
And 4, calculating the throat diameter of the engine in a working state through a throat diameter calculation module according to the average value of the throat temperature:
Dt＝18.00*(1+6e-6*(1612-25))＝18.23mm
and 5, calculating the mixing ratio through a mixing ratio calculation module according to the average flow values of the oxidant and the fuel:
rm＝93.2/56.6＝1.647
and 6, calculating the actual characteristic speed of the engine through an actual characteristic speed calculation module according to the throat diameter, the combustion chamber pressure, the oxidant and the fuel flow under the working state of the engine:
Figure BDA0001394160730000061
and 7, calculating the theoretical characteristic speed of the engine through a theoretical characteristic speed calculation module according to the mixing ratio of the engine and the pressure of the combustion chamber:
C*＝1742m/s
and 8, calculating the efficiency of the engine combustion chamber through an engine combustion chamber efficiency calculation module:
Figure BDA0001394160730000062
the foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes and modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention.

Claims (4)

1. A rocket engine combustion chamber efficiency analysis system is characterized by comprising
Data acquisition module for performing normal temperature T0Lower engine throat diameter Dt0Corresponding linear expansion coefficient alpha of combustion chamber material and engine combustion chamber pressure P during engine thermal testc(t), oxidizer and fuel mass flow rate qmo(t) and qmf(T) and throat temperature Tt(t) collecting;
the data average value calculation module is used for obtaining the average value of each parameter of the engine through a preset algorithm to obtain the average value of the pressure in the combustion chamber
Figure FDA0002992630140000011
Average mass flow of oxidant and fuel
Figure FDA0002992630140000012
And
Figure FDA0002992630140000013
mean value of throat temperature
Figure FDA0002992630140000014
A throat diameter calculation module for, according to a preset formula:
Figure FDA0002992630140000015
diameter D of throat portion of engine during operationtCalculating (1); wherein alpha is the linear expansion coefficient of the combustion chamber base material at the corresponding temperature;
a mixing ratio calculation module for, according to a preset formula: mixing ratio
Figure FDA0002992630140000016
Mixing ratio r of oxidant and fuel flowmCalculating (1);
actual characteristic velocityThe calculation module is used for calculating the following parameters according to a preset formula:
Figure FDA0002992630140000017
proceed to the actual characteristic speed C*Calculating (1); wherein D istIs the temperature corrected throat diameter;
theoretical characteristic velocity calculation module using combustion chamber pressure
Figure FDA0002992630140000018
And mixing ratio rmCalculating theoretical characteristic speed according to the principle of minimum Gibbs free energy
Figure FDA0002992630140000019
An engine combustion chamber efficiency calculation module for passing a predetermined formula
Figure FDA00029926301400000110
Efficiency η of combustion chamber of enginecAnd (4) calculating.
2. A rocket engine combustion chamber efficiency analysis method is characterized by comprising the following steps:
step 1, collecting normal temperature T through a data collection module0Lower engine throat diameter Dt0Inquiring the linear expansion coefficient alpha of the corresponding combustion chamber material;
step 2, carrying out hot test of the engine, and acquiring the pressure P of the combustion chamber of the engine through a data acquisition modulec(t), oxidizer and fuel mass flow rates qmo(t) and qmf(T), throat temperature Tt(t)；
Step 3, obtaining an arithmetic mean value of each parameter of the engine through a data mean value calculation module to obtain a mean value of the pressure in the combustion chamber
Figure FDA00029926301400000111
Average mass flow of oxidant and fuel
Figure FDA00029926301400000112
And
Figure FDA00029926301400000113
mean value of throat temperature
Figure FDA00029926301400000114
Step 4, calculating the throat diameter D of the engine during working through a throat diameter calculation module by utilizing the average value of the throat temperaturet：
Figure FDA0002992630140000021
Wherein alpha is the linear expansion coefficient of the combustion chamber base material at the corresponding temperature;
and calculating the mixing ratio r by a mixing ratio calculation module according to the average value of the mass flow of the oxidant and the fuelm：
Figure FDA0002992630140000022
Step 5, calculating the actual characteristic speed C through an actual characteristic speed calculation module*：
Figure FDA0002992630140000023
Wherein D istIs the temperature corrected throat diameter;
step 6, utilizing the pressure of the combustion chamber
Figure FDA0002992630140000024
And mixing ratio rmCalculating the theoretical characteristic speed according to the minimum Gibbs free energy principle by a theoretical characteristic speed calculation module
Figure FDA0002992630140000025
Step 7, calculating the efficiency eta of the engine combustion chamber through an engine combustion chamber efficiency calculation modulec：
Figure FDA0002992630140000026
3. A rocket engine combustion chamber efficiency analysis method according to claim 2, wherein when averaging each parameter, the average calculation interval is the same time interval in order to ensure that the parameters are taken from the same engine state.
4. A rocket engine combustion chamber efficiency analysis method according to claim 2, wherein when averaging parameters during engine commissioning, the length of commissioning procedure is required to ensure that there are sufficient stable sections of combustion chamber pressure, oxidant and fuel flow, throat temperature parameters, the stable section length being not less than 5 s.
CN201710766220.9A 2017-08-30 2017-08-30 Rocket engine combustion chamber efficiency analysis method and system Active CN108182297B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201710766220.9A CN108182297B (en) 2017-08-30 2017-08-30 Rocket engine combustion chamber efficiency analysis method and system

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201710766220.9A CN108182297B (en) 2017-08-30 2017-08-30 Rocket engine combustion chamber efficiency analysis method and system

Publications (2)

Publication Number Publication Date
CN108182297A CN108182297A (en) 2018-06-19
CN108182297B true CN108182297B (en) 2021-06-04

Family

ID=62545155

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201710766220.9A Active CN108182297B (en) 2017-08-30 2017-08-30 Rocket engine combustion chamber efficiency analysis method and system

Country Status (1)

Country Link
CN (1) CN108182297B (en)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115329695B (en) * 2022-10-11 2023-03-24 中国人民解放军63921部队 Rocket engine cooling channel outlet temperature estimation method and device

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103116705A (en) * 2013-02-06 2013-05-22 中国航天科技集团公司第六研究院第十一研究所 Fault simulated analysis method for afterburning cycle rocket engine
CN106134384B (en) * 2012-03-31 2014-06-18 上海空间推进研究所 The choosing method of rail control engine combustion chamber, a kind of space best features length

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6361629B2 (en) * 1998-11-12 2002-03-26 The United States Of America As Represented By The Secretary Of The Air Force Flowable solid propellant

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106134384B (en) * 2012-03-31 2014-06-18 上海空间推进研究所 The choosing method of rail control engine combustion chamber, a kind of space best features length
CN103116705A (en) * 2013-02-06 2013-05-22 中国航天科技集团公司第六研究院第十一研究所 Fault simulated analysis method for afterburning cycle rocket engine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
互击式喷嘴燃烧室燃烧效率实验;张蒙正 等;《推进技术》;20120229;第33卷(第1期);全文 *

Also Published As

Publication number Publication date
CN108182297A (en) 2018-06-19

Similar Documents

Publication Publication Date Title
CN112067304B (en) Method for measuring inlet flow of compressor in engine whole machine test
EP1930568A1 (en) Method and system for monitoring process states of an internal combustion engine
CN102213708A (en) Method for testing air leak rate of air preheater
CN113310536B (en) Method for measuring inlet flow of compressor in engine whole machine test
CN103775139A (en) Gap control system and method for turbine engine
CN110671219B (en) Method and system for controlling air-fuel ratio of gas engine
CN108182297B (en) Rocket engine combustion chamber efficiency analysis method and system
CN106441472A (en) Method for temperature drift inhibition of constant power type thermal type gas flowmeter
CN113361040B (en) Combustion chamber outlet temperature evaluation method under complete engine condition
CN109186815A (en) A kind of low temperature High Mach number detecting probe temperature calibration device
CN109948231B (en) Method and device for analyzing thermal cycle parameters of engine
US20160320061A1 (en) Online estimation of specific gravity of gas fuel
US20160131146A1 (en) Pressure sensor system for calculating compressor mass flow rate using sensors at plenum and compressor entrance plane
CN114910120B (en) Method for measuring high-speed airflow thermal parameters and flow parameters
CN109614722B (en) Modeling method for scroll engine full-state parameters based on fuzzy logic
JP2011047377A (en) Suction air amount measuring method and suction air amount measuring device
CN109655275B (en) Gas turbine cycle heat economy diagnosis method based on sensitivity coefficient method
US20170276622A1 (en) Method and system for gas temperature measurement
Buttsworth et al. Unsteady total temperature measurements downstream of a high-pressure turbine
CN110738011A (en) Temperature evaluation method and system for structural components in engines
CN111290319B (en) Water vapor preparation measuring device and system
CN110889217B (en) Correction method for outlet temperature of one-dimensional simulation model of combustion chamber of gas turbine
CN111473976B (en) Correction method for fuel flow of gas turbine engine
RU2389891C1 (en) Control method of air leakages and flow for turbine cooling in double-flow gas turbine engine
CN109520633A (en) A kind of double venturi air suction type measureing methods of high-temperature

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant