CN115329695B - Rocket engine cooling channel outlet temperature estimation method and device - Google Patents
Rocket engine cooling channel outlet temperature estimation method and device Download PDFInfo
- Publication number
- CN115329695B CN115329695B CN202211238640.7A CN202211238640A CN115329695B CN 115329695 B CN115329695 B CN 115329695B CN 202211238640 A CN202211238640 A CN 202211238640A CN 115329695 B CN115329695 B CN 115329695B
- Authority
- CN
- China
- Prior art keywords
- combustion chamber
- temperature
- operating condition
- substances
- formula
- 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
Links
Images
Classifications
-
- 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
- G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01K—MEASURING TEMPERATURE; MEASURING QUANTITY OF HEAT; THERMALLY-SENSITIVE ELEMENTS NOT OTHERWISE PROVIDED FOR
- G01K13/00—Thermometers specially adapted for specific purposes
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2113/00—Details relating to the application field
- G06F2113/08—Fluids
-
- 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
-
- 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
- G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
Landscapes
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- Theoretical Computer Science (AREA)
- Mathematical Physics (AREA)
- Fluid Mechanics (AREA)
- Mathematical Analysis (AREA)
- Mathematical Optimization (AREA)
- Computing Systems (AREA)
- Pure & Applied Mathematics (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- Geometry (AREA)
- General Engineering & Computer Science (AREA)
- Algebra (AREA)
- Testing Of Engines (AREA)
Abstract
The invention relates to a rocket engine cooling channel outlet temperature estimation method and a rocket engine cooling channel outlet temperature estimation device, and belongs to the technical field of liquid rocket engines. The method utilizes a formula based on the existing working condition parameters
Estimating the cooling passage outlet temperature under other conditions, wherein,T in in order to cool the temperature at the inlet of the channel,
for the coolant flow rate in the first operating condition,
for said first operating condition the combustion chamber propellant flow,p c0 the combustion chamber pressure is said first operating condition,T 0 the outlet temperature of the cooling channel is cooled for the first working condition,
for the coolant flow rate in the second condition,
for said second condition the combustion chamber propellant flow rate,p c1 and the pressure of the combustion chamber is the second working condition. The disclosed cooling channel exit temperature estimation model considers conventional tools such as coolant flow, thrust chamber gas flow, combustion chamber pressure, and the likeBesides condition parameters, the change of the geometric configuration of the thrust chamber can be considered, the scheme design of the engine system can be conveniently and rapidly developed, the state parameter distribution of the engine system can be obtained, and further more comprehensive system parameter optimization can be developed beneficially.
Description
Technical Field
The disclosure relates to a temperature estimation method and device, in particular to a rocket engine cooling channel outlet temperature estimation method and device, and belongs to the technical field of liquid rocket engines.
Background
By developing the scheme research of the liquid rocket engine system, the distribution of key node state parameters (including flow, pressure, temperature, power and the like) of the liquid rocket engine along the flow direction can be quickly obtained, on one hand, a student is helped to have popular and obvious deep knowledge on the engine, and on the other hand, the key node state parameters (including room pressure, mixing ratio, coolant flow and the like) are of great significance for optimizing and determining engine core parameters from the system level.
The scheme research of the liquid rocket engine system is developed, and a simple empirical model is generally adopted for quick estimation. For the power of a turbine pump and an electric pump, along-the-way, local pressure loss, flow and the like, mature mathematical estimation models are available. For liquid rocket engines employing regenerative cooling, since the fuel path must flow through the cooling passages, estimation of the cooling passage exit temperature becomes important in order to obtain engine state parameters at key nodes along the flow direction. Particularly for an expansion cycle liquid rocket engine, the temperature of the outlet of the cooling channel determines important state parameters such as power of a turbopump and the like, and directly influences the generation of the state parameters of the whole system scheme.
The current document (comparative study of a liquid oxygen methane expansion cycle variable thrust engine system scheme, proceedings of national defense science and technology university, 2020, 42 (3): 106-15.) provides a model of the change of the outlet temperature of a cooling channel along with working condition parameters under the condition that the geometric configuration of a thrust chamber is not changed, and the problem of estimation of the expansion cycle liquid rocket engine state parameters in the variable thrust process can be solved. According to the literature (scheme research on parallel electric heating cooperative pressure boosting variable thrust rocket engine [ J ]. Manned space, 2020, 26 (6): 702-9), by utilizing the model and the method, the system scheme research on the parallel electric heating cooperative pressure boosting variable thrust rocket engine is developed, but the influence of the change of the geometric configuration (the throat diameter of the thrust chamber, the outlet diameter of a spray pipe, the diameter of a combustion chamber and other straight sections) of the thrust chamber is not considered. Under the condition of the change of the geometric configuration of the thrust chamber, the outlet temperature of the cooling channel needs to be obtained through a complex method process, such as a cooling channel outlet temperature calculation method in the research of the regenerative cooling characteristic of the liquid oxymethane variable thrust rocket engine, or the outlet temperature of the cooling channel under a specific condition needs to be obtained through a simulation technology, and the efficiency of the cooling channel cannot meet the research requirement of a system. How to rapidly obtain the outlet temperature of the cooling channel under the condition of considering the influence of the geometric configuration of the thrust chamber based on working condition parameters (the inlet temperature of the cooling channel, the coolant flow, the propellant flow of the combustion chamber and the pressure of the combustion chamber) becomes a problem which is urgently needed to be solved and restricts the research of a liquid rocket engine system.
Disclosure of Invention
The purpose of the present disclosure is to overcome the drawbacks of the prior art and to solve the problem of fast estimation of the cooling passage exit temperature that can take into account the thrust chamber geometry effects, and to provide a rocket engine cooling passage exit temperature estimation method and apparatus.
The principle of the disclosure is based on a known set of working condition parameters, which can be obtained by the existing method, and the outlet temperature of the cooling channel of other working condition parameters under the same inlet temperature condition of the cooling channel is quickly obtained by a numerical calculation method, so that a rocket engine researcher can conveniently carry out system research.
The purpose of the present disclosure is achieved by the following technical solutions.
In a first aspect, the present disclosure provides a rocket engine cooling passage outlet temperature estimation method, comprising:
calculating a second operating condition cooling channel outlet temperature based on the first operating condition parameter by:
wherein the content of the first and second substances,T in in order to cool the temperature at the entrance of the channel,
based on the coolant flow in the first operating mode>
For said first operating condition the combustion chamber propellant flow,p c0 the combustion chamber pressure is said first operating condition,T 0 cooling a channel outlet temperature for the first operating condition>
Based on the coolant flow in the second operating mode>
For said second condition the combustion chamber propellant flow rate,p c1 combustion chamber pressure is said second operating condition.
In a second aspect, the present disclosure provides a rocket engine cooling passage exit temperature estimation device, a cooling passage exit temperature calculation module that calculates a second operating condition cooling passage exit temperature based on a first operating condition parameter by:
wherein the content of the first and second substances,T in in order to cool the temperature at the inlet of the channel,
based on the coolant flow in the first operating condition>
For said first operating condition the combustion chamber propellant flow,p c0 for the first operating condition the combustion chamber pressure,T 0 cooling the channel outlet temperature for the first operating condition->
Based on the coolant flow in the second operating mode>
For said second condition the combustion chamber propellant flow rate,p c1 and the pressure of the combustion chamber is the second working condition.
In a third aspect, the present disclosure provides an electronic device comprising:
at least one processor; and the number of the first and second groups,
a memory communicatively coupled to the at least one processor; wherein the memory stores instructions executable by the at least one processor to enable the at least one processor to perform the method of any of the embodiments of the first aspect.
In a fourth aspect, the present disclosure provides a computer-readable storage medium having stored thereon a computer program which, when executed by a processor, causes the processor to perform the method of any of the embodiments of the first aspect.
Advantageous effects
Compared with the prior art, the estimation model for the outlet temperature of the cooling channel is provided, the change of the geometric configuration of the thrust chamber can be considered besides the conventional working condition parameters such as coolant flow, gas flow of the thrust chamber and combustion chamber pressure, the scheme design of the engine system can be conveniently and rapidly developed, the distribution of the engine state parameters is obtained, and further more comprehensive optimization of the system parameters can be beneficially developed.
Drawings
FIG. 1 is a schematic view of a cooling channel;
FIG. 2 is a schematic view of a linearized thrust chamber profile;
FIG. 3 is a schematic diagram of the variation of cooling channel outlet temperature and relative error with chamber pressure calculated using the disclosed method and a prior art method;
fig. 4 is a schematic structural diagram of an electronic device according to an embodiment of the present disclosure.
Detailed Description
The present disclosure will be described in detail below with reference to specific embodiments shown in the drawings. These embodiments are not limited to the disclosure, and structural, methodological, or functional changes made by those of ordinary skill in the art in light of these embodiments are intended to be within the scope of the disclosure.
In the description of the present disclosure, it is to be understood that the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in the orientation or positional relationship indicated in the drawings for convenience in describing the present disclosure and for simplicity in description, and are not intended to indicate or imply that the referenced device or element must have a particular orientation, be constructed in a particular orientation, and be operated in a particular manner, and therefore should not be considered limiting to the disclosure. Furthermore, the terms "first", "second", etc. are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first," "second," etc. may explicitly or implicitly include one or more of that feature. In the description of the present disclosure, "a plurality" means two or more unless otherwise specified.
For the purpose of illustrating the objects, technical solutions and advantages of the embodiments of the present disclosure, the technical solutions in the embodiments of the present disclosure will be clearly and completely described below with reference to the drawings in the embodiments of the present disclosure.
When a researcher is researching a rocket engine system, the system performance of the researcher under different state parameters, such as the outlet temperature of a cooling channel, and the like, is required to be known frequently. Fig. 1 shows a schematic diagram of a cooling channel, and high-temperature and high-pressure fuel gas enters from a fuel gas inlet of a thrust chamber, is expanded to work and then is changed into high-speed, low-temperature and low-pressure fuel gas, and is discharged from a fuel gas outlet of the thrust chamber. The low-temperature coolant flows in from the coolant inlet (the temperature of the coolant inlet is consistent under different working conditions), the high-temperature coolant is changed after the thrust chamber is cooled, and the high-temperature coolant flows out from the coolant outlet (under different working conditions, the geometric configuration of the thrust chamber is different, or the geometric configuration of the same thrust chamber is the same but the pressure of different combustion chambers, and the temperature of the coolant is different when the coolant flows out from the outlet of the cooling channel). For this model, the system researchers need to know the temperature of coolant when flowing out from the export after getting into from the cooling channel entry, through with thrust room heat transfer process after to the thrust room cooling. This temperature determines important state parameters in rocket engine research such as turbopump power. Integrated Modeling and Analysis methods (Integrated Modeling and Analysis for LOX/Methane Expander Cycle Engine, AIAA 2006-4534) and liquid oxygen Methane variable thrust rocket Engine Regenerative Cooling characteristic research (liquid oxygen Methane variable thrust rocket Engine Regenerative Cooling characteristic research [ J ]. Manned space, 2022, 28 (1): 22-9.) are all related researches, but the methods are complicated, require complicated and complicated related system parameter setting, are not high, and cannot meet the system research requirements. The method in the scheme comparison research of the liquid oxygen methane expansion cycle variable thrust engine system can quickly estimate the outlet temperature of the cooling channel, but is only suitable for the situation that the geometric configuration of the thrust chamber is not changed. As shown in fig. 1, in the rocket engine system scheme research, the geometry of the thrust chamber is changed frequently, and the changed coolant outlet temperature is obtained through a complicated process again. Therefore, how to quickly obtain the coolant outlet temperature under the current working condition based on the existing working condition parameters considering the influence of the geometric configuration of the thrust chamber becomes a problem to be solved urgently in the research of rocket engines.
The invention provides a rocket engine cooling channel outlet temperature estimation method, which is based on a first working condition parameter and calculates a second working condition cooling channel outlet temperature according to the following formula:
wherein the content of the first and second substances,T in in order to cool the temperature at the entrance of the channel,
based on the coolant flow in the first operating mode>
For said first operating condition the combustion chamber propellant flow,p c0 for the first operating condition the combustion chamber pressure,T 0 cooling the channel outlet temperature for the first operating condition->
Based on the coolant flow in the second operating mode>
For said second condition the combustion chamber propellant flow rate,p c1 and the pressure of the combustion chamber is the second working condition.
By using the above formula, the outlet temperature of the cooling channel under other working conditions can be quickly obtained based on the parameters under one working condition, and the requirement for obtaining the outlet temperature of the coolant in time under the condition of variable parameters in the research of the rocket engine is met. Also, the above estimation method is applicable to a case where the geometry of the thrust chamber changes.
The above estimation method modeling process is detailed below:
for the cooling channel shown in fig. 1, neglecting wall surface heat conduction, according to the equal side flow of the gas side and the coolant side, it can be obtained:
wherein the content of the first and second substances,
based on the coolant flow>
Is the constant-pressure specific heat capacity of the coolant,h g the convective heat transfer coefficient is measured for the gas,T c is the temperature of the combustion chamber, and,T aw is the temperature of the side wall surface of the fuel gas,Dthe diameter of the combustion chamber is taken as the diameter,Lthe axial length of the combustion chamber is taken as the length,Tit is the temperature that is set for the purpose,yare axis direction coordinates.
Applying the integral median theorem to the left side of the formula (1) willh g The concrete expression (refer to formula 4-11 on page 104 of 'liquid rocket engine modern engineering design') is substituted into the right side of formula (1), and the following can be obtained:
wherein the content of the first and second substances,
is the integrated median value from the coolant side inlet to the outlet, based on the measured value>
Is the gas flow of the thrust chamber>
To cool the channel temperature rise.
According to the literature (Jingpeng, li Qing Lian, chengpo, zhang Beichen, liquid oxygen methane expansion cycle variable thrust engine system scheme contrast research [ J)]The university of defense science and technology, 2020, 42 (3): 106-15.), neglecting the left-side median integral of formula (2)
Temperature difference between the gas side and the right side>
The following results were obtained:
wherein the content of the first and second substances,kis a constant.
To further simplify the derivation, it is assumed that the thrust chamber profile varies linearly, as shown in FIG. 2.
The thrust chamber profile expression is then:
wherein the content of the first and second substances,a、b、c、dis a coefficient of diameter of a straight section of the combustion chamberD c Diameter of throatD t Diameter of the nozzle outletD e Length of equal straight sectiony 1 Length of the constrictiony 2 Length of the expanded sectiony 3 And (4) correlating.
Substituting the linear expression into the integral term of the formula (3) can obtain:
assuming a characteristic length of the combustion chamberL c In the general context (references Mazda, lijiawen, tianemai, zhangui. Liquid rocket engine design [ M)]Beijing university of aerospace Press 2011.) is variable, the combustion chamber volumeV c =L c ·A t Variable, can make:
y 1 =(D c /2)·k 1 （6）
that is, the length of the equal straight section of the combustion chamber changes proportionally with the diameter of the combustion chamber. Wherein, the first and the second end of the pipe are connected with each other,A t is the area of the throat part,k 1 is a constant.
(this section demonstrates:
assume another expression for the combustion chamber volume:
wherein the content of the first and second substances,
assumed shrinkage ratio
Is constant and at the same time willy 1 Substituting the expression into the above equation, then:
V c = k 2 ·A t 1.5 ；
thus, whenL c Is changed, thenA t There is a solution. )
A mixture of (6), (7) and
substitution into formula (5) can yield:
wherein the content of the first and second substances,a、care respectively asα、βIs a constant.
Ignore
Then equation (8) becomes:
thus, equation (3) becomes:
as can be seen from the expression, the cooling channel temperature rise
Along with the throat area of the thrust chamber>
May vary.
The flow formula of the combustion chamber is
Wherein the content of the first and second substances,
is the specific heat ratioγAs a function of (c).
Neglecting the specific heat ratio of the fuel gas when the mixing ratio is not changedγGas constant ofRTemperature of combustion chamberT c When the formula (11) is substituted into the formula (10),
wherein the content of the first and second substances,k 2 ，k 3 ，k 4 ，k 5 is a constant.
At given reference condition parameters
、/>
Andp c0 after, the outlet temperature of the cooling channel under the reference working conditionT 0 The expression is as follows,
a parameter of a certain operating condition
、/>
Andp c1 lower cooling channel exit temperatureT 1 Expression is as>
Eliminating the constants in equation (14) using equation (13)k 5 Then obtainT 1 Is expressed in the final expression of (a),
wherein the content of the first and second substances,T in to cool the channel inlet temperature (consistent across different operating conditions).
From the above, equation 10 fully considers the influence of the geometry of the thrust chamber on the outlet temperature of the cooling channel, so that finally the estimation of the outlet temperature of the cooling channel in the variable geometry or the constant geometry of the thrust chamber can be obtained based on the reference working condition by equation 15.
Experimental verification
The first verification method comprises the following steps:
1. the idea in the literature (Study on the heat transfer characteristics of regenerative cooling for LOX/LCH4 variable cooling engine [ J ]. Case students in Thermal Engineering, 2021, 28: 1-12. And research on regenerative cooling characteristics of liquid oxygen methane variable thrust rocket engine [ J ]. Manned space, 2022, 28 (1): 22-9.) was used to calculate a coolant flow of 2.02kg/s, a combustion chamber propellant flow of 8.48kg/s, and a cooling channel outlet temperature of 291K under a chamber pressure of 4.5MPa, as a first condition parameter.
2. The working conditions in the following table were calculated using the calculation formula of the method of the present disclosure, and the calculation results of the outlet temperature of the cooling channel and the error are shown in fig. 3. It can be seen that there is some error between the calculation results using the method of the present invention and the calculation results of the literature methods (Study on the heat transfer characteristics of regenerative cooling for LOX/LCH4 variable cooling rock engine [ J ]. Case students in Thermal Engineering, 2021, 28: 1-12. And liquid oxygen methane thrust rocket engine regenerative cooling characteristics Study [ J ]. Manned space, 2022, 28 (1): 22-9.), and the maximum error is substantially less than 10%. Considering that the method and the experimental result in the literature have certain errors, and the literature (research on liquid nitrogen flowing and heat exchange characteristics in microchannels, 2007, shanghai university of transportation, doctor's academic thesis, page 66) indicates that the error of the calculation result obtained by the theoretical convective heat exchange coefficient correlation formula adopted by the simulation is reasonable within 30%, so that the method disclosed by the invention is reliable.
TABLE 1 different working Condition parameters
| Serial number | Coolant flow/(kg/s) | Propellant flow/(kg/s) | Room pressure/ |
| 1. | 1.62 | 6.78 | 3.6 |
| 2. | 1.21 | 5.09 | 2.7 |
| 3. | 1.01 | 4.24 | 2.25 |
| 4. | 1.35 | 5.67 | 3 |
| 5. | 1.08 | 4.54 | 2.4 |
| 6. | 0.94 | 3.95 | 2.1 |
| 7. | 0.67 | 2.81 | 1.5 |
| 8. | 0.54 | 2.27 | 1.2 |
And a second verification method:
1. review of literature: an Integrated Modeling and Analysis method (Integrated Modeling and Analysis for a LOX/Methane Expander Cycle Engine, AIAA 2006-4534) of a liquid oxygen Methane expansion Cycle Engine determines a coolant inlet temperature of 116.7K, a coolant flow of 6.865kg/s, a combustion chamber propellant flow of 30.917kg/s, a combustion chamber pressure of 5.861MPa and a Cooling channel outlet temperature of 526.2K.
2. The calculation formula of the method disclosed by the invention is utilized to calculate the coolant flow rate to be 6.057kg/s, the propellant flow rate to be 27.28kg/s and the outlet temperature of the cooling channel under the working condition of 5.171MPa of the combustion chamber to be 536.6K, the literature query result is 543.9K, and the error is 1.34%, so that the method disclosed by the invention is effective.
The present disclosure also provides a rocket engine cooling passage outlet temperature estimation device, including a cooling passage outlet temperature calculation module that calculates a second operating condition cooling passage outlet temperature based on a first operating condition parameter by:
wherein the content of the first and second substances,T in in order to cool the temperature at the inlet of the channel,
based on the coolant flow in the first operating mode>
For said first operating condition the combustion chamber propellant flow,p c0 for the first operating condition the combustion chamber pressure,T 0 cooling the channel outlet temperature for the first operating condition->
Based on the coolant flow in the second operating mode>
For said second condition the combustion chamber propellant flow rate,p c1 and the pressure of the combustion chamber is the second working condition.
Alternatively to this, the first and second parts may,T 0 based on
、/>
、p c0 The method is obtained by a cooling channel outlet temperature calculation method in the research of the regenerative cooling characteristic of the liquid oxygen methane variable thrust rocket engine.
Optionally, the first operating condition parameter is obtained by consulting integrated modeling and analysis method data of the liquid oxymethane expansion cycle engine.
For the apparatus embodiment, since it is substantially similar to the method embodiment, the description is relatively simple, and reference may be made to the partial description of the method embodiment for relevant points.
Fig. 4 is a schematic structural diagram of an electronic device according to an embodiment of the present disclosure, where the electronic device may execute the processing flow provided by the foregoing method embodiment, and as shown in fig. 4, the electronic device 110 includes: memory 111, processor 112, computer programs, and communications interface 113; wherein the computer program is stored in the memory 111 and is configured to be executed by the processor 112 for performing the method as described above.
In addition, the embodiment of the present disclosure also provides a computer readable storage medium, on which a computer program is stored, the computer program being executed by a processor to implement the method of the above embodiment. Those of ordinary skill in the art will understand that: all or a portion of the steps of implementing the above-described method embodiments may be performed by hardware associated with program instructions. The foregoing program may be stored in a computer-readable storage medium. When executed, the program performs steps comprising the method embodiments described above; and the aforementioned storage medium includes: various media that can store program codes, such as ROM, RAM, magnetic or optical disks.
Specific examples are given in this specification for the purpose of illustrating the disclosure and implementations. The details introduced in the examples are not intended to limit the scope of the claims but rather to aid in understanding the present disclosure. Those skilled in the art will understand that: although the description is given in terms of embodiments, not every embodiment includes only a single embodiment, and such description is for clarity only, and those skilled in the art will recognize that the embodiments described herein may be combined as a whole to form other embodiments as would be understood by those skilled in the art. And that various modifications, changes, or alterations to the steps of the preferred embodiments are possible without departing from the spirit and scope of this disclosure and the appended claims. Therefore, the disclosure should not be limited to the disclosure of the preferred embodiments and drawings.
Claims (8)
1. A rocket engine cooling passage outlet temperature estimation method is characterized by comprising the following steps: calculating a second operating condition cooling passageway exit temperature based on the first operating condition parameter by:
wherein the content of the first and second substances,T in in order to cool the temperature at the inlet of the channel,
for the coolant flow in the first condition,
for the first operating condition the combustion chamber propellant flow,p c0 for the first operating condition the combustion chamber pressure,T 0 the channel exit temperature is cooled for the first operating condition,
for the coolant flow rate in the second condition,
for the second operating mode the flow of propellant in the combustion chamber,p c1 the pressure of the combustion chamber is the second working condition;
the modeling process of the above formula is:
neglecting wall heat conduction, according to the equal side heat flow of gas and coolant, can obtain:
wherein the content of the first and second substances,
as the flow rate of the coolant, it is,
is the constant-pressure specific heat capacity of the coolant,h g the convective heat transfer coefficient is measured for the gas,T c is the temperature of the combustion chamber, and,T aw is the temperature of the side wall surface of the fuel gas,Dthe diameter of the combustion chamber is taken as the diameter,Lis burnedThe axial length of the burning chamber is long,Tis the temperature of the liquid to be treated,yis an axis direction coordinate;
applying the integral median theorem to the left side of the formula (1) willh g The specific expression is substituted into the right side of the formula (1), and the following can be obtained:
wherein the content of the first and second substances,
is the integrated median value of the coolant side inlet to outlet,
is the flow rate of the gas in the thrust chamber,
raising the temperature of the cooling channel;
neglecting the left-hand median integral of equation (2)
Temperature difference of right gas side
The following results were obtained:
wherein the content of the first and second substances,kis a constant;
if the thrust chamber profile changes linearly, the thrust chamber profile expression is:
wherein the content of the first and second substances,a、b、c、dis a coefficient of diameter of a straight section equal to that of the combustion chamberD c Diameter of throatD t Diameter of the nozzle outletD e Length of equal straight sectiony 1 Length of the constriction sectiony 2 Length of the expanded sectiony 3 Correlation;
substituting the profile expression into the integral term of formula (3) to obtain:
setting characteristic length of combustion chamberL c Variable in the general range, the combustion chamber volumeV c =L c ·A t Variable, can make:
y 1 =(D c /2)·k 1 （6）
that is, the length of the equal straight section of the combustion chamber changes with the diameter of the combustion chamber in equal proportion; wherein the content of the first and second substances,A t is the area of the throat part,k 1 is a constant;
a general formula (6),
And
substitution into formula (5) can give:
neglect of
Then equation (8) becomes:
thus, equation (3) becomes:
as can be seen from the expression, the cooling channel temperature rise
Throat area of thrust chamberA t Is changed;
the flow formula of the combustion chamber is
Wherein the content of the first and second substances,
is the specific heat ratioγA function of (a);
neglecting the specific heat ratio of the fuel gas when the mixing ratio is not changedγGas constant ofRTemperature of combustion chamberT c When the formula (11) is substituted into the formula (10),
wherein the content of the first and second substances,k 2 ，k 3 ，k 4 ，k 5 is a constant;
at a given first operating condition parameter
、
Andp c0 then, the outlet temperature of the cooling passage under the first working conditionT 0 The expression is as follows,
second operating condition parameter
、
Andp c1 lower cooling channel exit temperatureT 1 The expression is as follows,
eliminating the constants in equation (14) using equation (13)k 5 Then obtainT 1 Is expressed in the final expression of (a),
wherein the content of the first and second substances,T in to cool the channel inlet temperature, it is consistent under different conditions.
2. A rocket engine cooling passage exit temperature estimation method as recited in claim 1, wherein:T 0 based on
、
、p c0 Passing liquidAnd the calculation method of the outlet temperature of the cooling channel in the research of the regenerative cooling characteristic of the oxymethane variable thrust rocket engine is obtained.
3. A rocket engine cooling passage exit temperature estimation method as recited in claim 1, wherein: the first operating condition parameter is obtained by consulting data of an integrated modeling and analysis method of the liquid oxygen methane expansion cycle engine.
4. A rocket engine cooling passage outlet temperature estimation device, characterized by: a cooling passage outlet temperature calculation module is included that calculates a second operating condition cooling passage outlet temperature based on the first operating condition parameter by:
wherein the content of the first and second substances,T in in order to cool the temperature at the inlet of the channel,
for the coolant flow in the first condition,
for the first operating condition the combustion chamber propellant flow,p c0 for the first operating condition the combustion chamber pressure,T 0 the channel exit temperature is cooled for the first operating condition,
for the coolant flow rate in the second condition,
for the second operating mode the flow of propellant in the combustion chamber,p c1 the pressure of the combustion chamber is in a second working condition;
the modeling process of the above formula is:
neglecting wall heat conduction, according to the equal side heat flow of gas and coolant, can obtain:
wherein, the first and the second end of the pipe are connected with each other,
as the flow rate of the coolant, it is,
is the constant-pressure specific heat capacity of the coolant,h g the convective heat transfer coefficient is measured for the gas,T c is the temperature of the combustion chamber, and,T aw is the temperature of the side wall surface of the fuel gas,Dthe diameter of the combustion chamber is taken as the diameter,Lthe axial length of the combustion chamber is,Tit is the temperature that is set for the purpose,yis an axis direction coordinate;
applying integral median theorem to the left side of equation (1) willh g The specific expression is substituted into the right side of the formula (1), and the following can be obtained:
wherein the content of the first and second substances,
is the integrated median value of the coolant side inlet to outlet,
is the flow rate of the gas in the thrust chamber,
raising the temperature of the cooling channel;
neglecting the left-hand median integral of equation (2)
Right side, right sideGas side temperature difference
The following results were obtained:
wherein the content of the first and second substances,kis a constant;
if the thrust chamber profile changes linearly, the thrust chamber profile expression is:
wherein, the first and the second end of the pipe are connected with each other,a、b、c、dis a coefficient of diameter of a straight section of the combustion chamberD c Diameter of throatD t Diameter of the nozzle outletD e Length of equal straight sectiony 1 Length of the constrictiony 2 Length of the expanded sectiony 3 Correlation;
substituting the linear expression into the integral term of the formula (3) can obtain:
setting characteristic length of combustion chamberL c Variable in the general range, the combustion chamber volumeV c =L c ·A t Variable, can make:
y 1 =(D c /2)·k 1 （6）
that is, the length of the equal straight section of the combustion chamber changes with the diameter of the combustion chamber in equal proportion; wherein the content of the first and second substances,A t is the area of the throat part,k 1 is a constant;
general formula (II)（6）、
And
substitution into formula (5) can yield:
neglect of
Then equation (8) becomes:
thus, equation (3) becomes:
as can be seen from the expression, the cooling channel temperature rise
Throat area of thrust chamberA t Is changed;
the flow formula of the combustion chamber is
Wherein the content of the first and second substances,
is the specific heat ratioγA function of (a);
neglecting the specific heat ratio of the fuel gas when the mixing ratio is not changedγGas constant ofRTemperature of combustion chamberT c When the formula (11) is substituted into the formula (10),
wherein the content of the first and second substances,k 2 ，k 3 ，k 4 ，k 5 is a constant;
at a given first operating condition parameter
、
Andp c0 then, the outlet temperature of the cooling passage under the first working conditionT 0 The expression is as follows,
second operating condition parameter
、
Andp c1 lower cooling channel exit temperatureT 1 The expression is as follows,
eliminating the constants in equation (14) using equation (13)k 5 Then obtainT 1 Is expressed in the final expression of (a),
wherein the content of the first and second substances,T in to cool the channel inlet temperature, it is consistent under different conditions.
5. A rocket engine cooling passage outlet temperature estimating device according to claim 4, wherein:T 0 based on the
、
、p c0 The method is obtained by a cooling channel outlet temperature calculation method in the research of the regenerative cooling characteristic of the liquid oxygen methane variable thrust rocket engine.
6. The rocket engine cooling passage outlet temperature estimation device of claim 4, wherein: the first operating condition parameter is obtained by consulting data of an integrated modeling and analysis method of the liquid oxygen methane expansion cycle engine.
7. An electronic device, characterized in that the electronic device comprises:
at least one processor; and the number of the first and second groups,
a memory communicatively coupled to the at least one processor; wherein the memory stores instructions executable by the at least one processor to enable the at least one processor to perform the method of any one of claims 1-3.
8. A computer-readable storage medium, on which a computer program is stored, which program, when being executed by a processor, causes the processor to carry out the method of any one of claims 1-3.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211238640.7A CN115329695B (en) | 2022-10-11 | 2022-10-11 | Rocket engine cooling channel outlet temperature estimation method and device |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211238640.7A CN115329695B (en) | 2022-10-11 | 2022-10-11 | Rocket engine cooling channel outlet temperature estimation method and device |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN115329695A CN115329695A (en) | 2022-11-11 |
| CN115329695B true CN115329695B (en) | 2023-03-24 |
Family
ID=83915163
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211238640.7A Active CN115329695B (en) | 2022-10-11 | 2022-10-11 | Rocket engine cooling channel outlet temperature estimation method and device |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115329695B (en) |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10527003B1 (en) * | 2015-04-12 | 2020-01-07 | Rocket Lab Usa, Inc. | Rocket engine thrust chamber, injector, and turbopump |
| CN108182297B (en) * | 2017-08-30 | 2021-06-04 | 上海空间推进研究所 | Rocket engine combustion chamber efficiency analysis method and system |
| CN111950148B (en) * | 2020-08-11 | 2023-10-03 | 江苏深蓝航天有限公司 | Method and equipment for calculating temperature of inner wall of liquid rocket thrust chamber test run |
| CN112832929B (en) * | 2021-03-05 | 2022-05-24 | 中国科学院力学研究所 | Method for designing cooling structure for equal inner wall surface temperature of rocket engine |
| CN112948982B (en) * | 2021-04-14 | 2022-11-01 | 北京航空航天大学 | Reliability analysis method and device for reusable rocket engine thrust chamber |
-
2022
- 2022-10-11 CN CN202211238640.7A patent/CN115329695B/en active Active
Non-Patent Citations (2)
| Title |
|---|
| Study on the heat transfer characteristics of regenerative cooling for LOX/LCH4 variable thrust rocket engine;Jie Song et al.;《Case Studies in Thermal Engineering》;20211215;全文 * |
| 液氧甲烷变推力火箭发动机再生冷却特性研究;孙郡 等;《载人航天》;20220215;第28卷(第1期);全文 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN115329695A (en) | 2022-11-11 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Oliet et al. | Parametric studies on automotive radiators | |
| Zuo et al. | Performance evaluation of regenerative cooling/film cooling for hydrocarbon fueled scramjet engine | |
| CN110968967B (en) | Heat transfer coupling simulation order reduction method for underground pipe heat exchanger | |
| Carcasci et al. | A numerical procedure to design internal cooling of gas turbine stator blades | |
| Li et al. | A Unfied computational formulation for multi-component and multi-phase flows | |
| Xie et al. | A method of rib-bed plate enhancing heat transfer in hydrogen rocket engine chamber wall | |
| CN111144054A (en) | Modeling method for natural circulation characteristic of villiaumite cooling high-temperature reactor passive waste heat discharge system | |
| Saviers et al. | Design and validation of topology optimized heat exchangers | |
| Agaie et al. | A novel technique of reduce order modelling without static correction for transient flow of non-isothermal hydrogen-natural gas mixture | |
| CN114154432B (en) | Printed circuit board type heat exchanger calculation method based on node division method | |
| CN115329695B (en) | Rocket engine cooling channel outlet temperature estimation method and device | |
| Pietropaoli et al. | Design for additive manufacturing: Internal channel optimization | |
| Shan et al. | Numerical modeling on pumping performance of piccolo-tube multi-nozzles supersonic ejector in an oil radiator passage | |
| CN114385960A (en) | Energy average temperature-based dividing wall type heat exchanger performance calculation method | |
| Jokhakar et al. | A CFD-RTE model for thermal analysis of regeneratively cooled rocket engines | |
| Naraghi et al. | A simple approach for thermal analysis of regenerative cooling of Rocket engines | |
| Yang et al. | A fin analogy model for thermal analysis of regeneratively cooled rocket engines | |
| Daimon et al. | Film cooling performance analysis of a full-scale liquid rocket engine combustion chamber based on a coupled combustion and heat transfer simulation | |
| Yan et al. | Retrofit design of composite cooling structure of a turbine blade by fluid networks and conjugate heat transfer methods | |
| Bedick et al. | Validation and Application of a Reduced-Order Rotating Detonation Engine Inlet and Fill Zone Model | |
| Chatzi et al. | Metal foam recuperators on micro gas turbines: Multi-objective optimisation of efficiency, power and weight | |
| CN115758627A (en) | Variable working condition parameter estimation method and device for open expansion cycle engine | |
| Leonardi et al. | A zooming approach to investigate heat transfer in liquid rocket engines with ESPSS propulsion simulation tool | |
| Wei et al. | A simplified numerical method based on anisotropic porous media model for the precooler of the hypersonic precooled engine | |
| Guo et al. | Research on a new method to optimize the thermal characteristics of an engine nacelle cooling module |
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 |