CN115982943A - Volumetric dynamics one-dimensional modeling method for scramjet engine - Google Patents
Volumetric dynamics one-dimensional modeling method for scramjet engine Download PDFInfo
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Abstract
The invention discloses a volumetric dynamics one-dimensional modeling method of a scramjet engine, which comprises the following steps of: step 1), establishing mathematical models of an air inlet channel, a combustion chamber and a tail nozzle of the scramjet engine; step 2) establishing a scramjet engine scramjet mode without shock wave mathematical models of the deflagration oblique shock wave mode and the subfiring mode; and step 3), establishing a dynamic model of the scramjet engine according to the volumetric dynamics. The method utilizes the volume dynamics to establish the scramjet engine one-dimensional model, can quickly calculate the change of the axial flow field parameters of the combustion chamber, and can meet the design requirement of an engine control system.
Description
Technical Field
The invention belongs to the technical field of scramjet control, and particularly relates to a scramjet volumetric dynamics one-dimensional modeling method.
Background
The basic and important point of engine control system research is engine model, and the research on establishing scramjet engine mathematical model starts from the proposal of supersonic combustion concept, including the research on various supersonic flow and physical effect modeling in combustion and engine complete machine modeling. Heiser and Pratt establish a scramjet model HAP. In the calculation of the combustion chamber, the model assumes an ideal chemical balance, adopts a rational equation to fit a heat release rule, describes the flow process inside the combustion chamber by dividing the combustion chamber into three heat transfer processes of adiabatic compression, isobaric heat release and expansion heat release, and simplifies the analysis process. However, due to the assumption of isobaric combustion, fuel single injection and the empirically determined onset of heat release and flow reattachment points, the model is not accurate enough at high mach numbers.
Many researchers in China also put forward a plurality of one-dimensional models of scramjet engines. The Caocheng peak of Hagong analyzes the combustion mode conversion boundary of the scramjet engine, researches the interaction between the combustion chamber and the isolation section, and establishes a one-dimensional steady-state mathematical model of the combustion chamber. The existing domestic and foreign documents mainly aim at the research of a stable state model of the scramjet, the research of a dynamic model of the scramjet designed aiming at a control system is less, only a few documents research a dynamic real-time model of an engine component level, such as Liuming epitaxy and Zhang Hai wave of south navigation, and a combustion chamber volume effect is adopted, so that a real-time model of the scramjet component level is established, and a closed-loop simulation test is carried out. Although the component-level dynamic model of the scramjet can describe the dynamic characteristics of the engine, the component-level dynamic model cannot be used for controlling the distribution parameters of the scramjet, so that a one-dimensional dynamic model of the scramjet needs to be established.
Disclosure of Invention
The invention aims to: in order to overcome the defects in the prior art, the invention provides a volumetric dynamics one-dimensional modeling method for a scramjet engine, which comprises the steps of establishing a mathematical model of an air inlet channel, a combustion chamber and a tail nozzle of the scramjet engine and a mathematical model of a scramjet engine in a scramjet mode, wherein the mathematical model is in a scramjet mode and is in a scramjet mode; and establishing a dynamic model of the scramjet engine according to the volumetric dynamics on the basis. The method utilizes the volume dynamics to establish the scramjet engine one-dimensional model, can quickly calculate the change of the axial flow field parameters of the combustion chamber, and can meet the design requirement of an engine control system.
The technical scheme is as follows: in order to achieve the purpose, the scramjet engine volume dynamics one-dimensional modeling method comprises the following steps:
step 1: establishing mathematical models of an air inlet channel, a combustion chamber and a tail nozzle of the scramjet engine according to the working principle of each part of the scramjet engine;
step 2: according to the combustion field mechanism of the scramjet engine, establishing a scramjet engine with scramjet mode mathematical models of the deflagration oblique shock wave mode and the subfiring mode;
and 3, step 3: and (3) establishing a dynamic model of the scramjet based on the modal model of the scramjet established in the step (2) according to the volume dynamics principle.
Further, the specific steps in step 1 are as follows:
step 1-1: based on experimental data, a polynomial fitting method is adopted to give total temperature distribution along the whole flow direction
Wherein x is the axial position coordinate (m) of the combustion chamber of the engine, T t (x) Is the total temperature at the x position, T t2 For total insulation section inlet temperature, τ is total heating ratio, θ is heat release rate, and is an empirical constant from 1 to 10, χ: the non-dimensional axial position of the shaft,
x i is the fuel injection point position, x 4 Is the combustion chamber exit location.
Step 1-2: coefficient of friction prediction using the formula
In the formula, C f : local friction coefficient, k: gas adiabatic index, re x : the number of reynolds numbers in the local area,
rho is the local density, V is the local speed, x is the local axial position coordinate with the engine inlet as the origin, mu is the local aerodynamic viscosity, and the local aerodynamic viscosity is calculated by the Sutherland formula
In the formula, mu 0 Is the viscosity coefficient under standard atmospheric conditions, T s Is the Soxhlet constant, dependent on the nature of the gas, T c =273.16k, t is the local temperature.
Step 1-3: deducing to obtain a one-dimensional control equation, mach number Ma and total pressure p of a flow field in a combustion chamber of the scramjet engine from a basic control equation set of gas dynamics t The variation relationship with the one-dimensional coordinate x is respectively
In the formula, A: cross-sectional area of combustion chamber, T t : total temperature; d: hydraulic diameter of the combustion chamber;
step 1-4: on the basis of obtaining the Mach number and the total pressure, other parameters are calculated by a gas dynamic function;
in the formula, T: static temperature, P: static pressure, V: speed, R: molar gas constant.
Further, the specific steps of step 2 are as follows:
step 2-1: establishing a conversion boundary between different combustion modes, wherein the conversion boundary between the scramjet-free mode and the scramjet mode is as follows: when the lowest Mach number of the combustion chamber is larger than 0.762 times of the Mach number of the inlet of the isolation section, the combustion chamber works in a super-combustion shock-wave-free mode, and on the contrary, the combustion chamber works in a super-combustion oblique shock-wave mode; the transition boundary between the deflagration oblique shock mode and the sub-combustion mode is: when the lowest Mach number of the combustion chamber is more than or equal to 1, the combustion chamber works in a super-combustion oblique shock wave mode; on the contrary, the combustion chamber works in a sub-combustion mode;
step 2-2: when the scramjet engine is in a scramjet mode, the whole flow channel flows at supersonic speed, and the separation of the isolation section and the combustion chamber does not generate boundary layer separation; taking the isolation section as a one-dimensional steady fanno flow section, taking the geometric profile of the engine as the flow profile of the variable cross-section heating pipe flow of the combustion chamber, and solving an ordinary differential equation set to obtain the distribution of each pneumatic parameter of the isolation section and the combustion chamber;
step 2-3: when the scramjet engine is in a scramjet mode, the whole flow channel flows at an ultrasonic speed, and the isolation section and the combustion chamber generate boundary layer separation; the front section of the isolation section is a Vanuno flow section, the rear section of the isolation section is a shock wave compression section, the front section of the combustion chamber is a combustion separation section, and the rear section of the combustion chamber is a separation attachment section; the isolation section shock wave compression section and the combustion chamber separation section take the geometric profile of the engine minus the separation area as a flow profile, and the isolation section fanno flow section and the combustion chamber separation attachment section take the geometric profile of the engine as the flow profile;
step 2-4: when the scramjet engine is in a sub-combustion mode, a strong shock wave string can appear in the isolation section due to high back pressure generated by the combustion chamber, the front section of the isolation section is still a fanno flow section, and the rear section of the isolation section is still a shock wave compression section; the combustion chamber inlet is subsonic, no boundary layer separation exists, and the geometric profile of the engine is used as a flow profile.
Further, the specific steps in step 3 are as follows:
step 3-1: taking the inlet of the combustion chamber as the inlet of the cavity, taking the position of the required point as the outlet of the cavity, and carrying out volume dynamics analysis after determining the position of the outlet of the cavity;
step 3-2: establishing a linear ordinary differential equation of the temperature and the pressure in the cavity to obtain an update equation of the temperature and the pressure based on the time step length:
in the formula, the subscript in: chamber inlet parameters, subscript out: outlet parameter of the chamber, V C : the volume of the containing cavity body is increased,
eta: combustion efficiency, W: flow rate, W f : fuel flow rate H f : fuel calorific value, h: total enthalpy, h C : a static enthalpy of fuel;
step 3-3: taking the steady state result as an initial value, calculating the total temperature and the total pressure by the formulas (9) and (10), and calculating a Mach number by a flow formula to obtain the static temperature and the static pressure;
step 3-4: and (3) solving by adopting the modal modeling method in the step (2) on the basis of obtaining the parameters, so that all the parameters of the flow field at the time t are obtained and serve as initial values at the time t +1, and the calculation is carried out in a circulating mode until the total temperature and the total pressure change rate are 0.
Has the advantages that: compared with the prior art, the volume dynamics one-dimensional modeling method for the scramjet engine, which is provided by the invention, has the following technical effects:
(1) According to the invention, through the analysis of the mechanism of the scramjet engine, the multidimensional flow is reduced to one-dimensional flow along the axial direction of the engine, and the established model can be well matched with experimental data, so that the real-time property of the model is improved;
(2) The method adopts a volume dynamics method to establish the one-dimensional dynamic model of the scramjet, can reflect the dynamic working characteristics of the scramjet, and can meet the design requirements of an engine control system.
Drawings
FIG. 1 is a schematic diagram of a scramjet engine.
FIG. 2 is a schematic view of the internal volume effect of the combustion chamber volume.
Fig. 3 is a flow chart of the present invention.
FIG. 4 is a graph of model static pressure output versus experimental data.
Figure 5 is a graph of model mach number output versus experimental data.
Fig. 6 is a tracking diagram of the wide variation of the thrust of the PI controller under H =27km and ma =6 flight conditions.
Fig. 7 is a graph of the change of the output fuel flow of the PI controller under H =27km and ma =6 flight conditions.
Detailed Description
The technical solutions of the present invention will be described in further detail with reference to the accompanying drawings, and it should be understood that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
In order to overcome the defects in the prior art, the invention provides a scramjet engine volume dynamics one-dimensional modeling method, as shown in fig. 3, according to the working principle of each part of the scramjet engine, a mathematical model of an air inlet channel, a combustion chamber and a tail nozzle of the scramjet engine and a mathematical model of a scramjet engine in a scramjet engine scramjet mode, a scramjet engine in a scramjet engine scramjet mode and a mathematical model in a scramjet engine in a scramjet mode and a scramjet engine in a scramjet mode are established; and establishing a dynamic model of the scramjet engine according to the volumetric dynamics on the basis. The method utilizes the volume dynamics to establish the scramjet engine one-dimensional model, can quickly calculate the change of the axial flow field parameters of the combustion chamber, and can meet the design requirement of an engine control system.
A scramjet engine volume dynamics one-dimensional modeling method comprises the following steps:
step 1: establishing mathematical models of an air inlet channel, a combustion chamber and a tail nozzle of the scramjet engine according to the working principle of each part of the scramjet engine;
step 2: according to the combustion field mechanism of the scramjet engine, establishing a mathematical model of a scramjet engine scramjet mode, a scramjet mode and a substring mode;
and 3, step 3: and (3) establishing a dynamic model of the scramjet based on the modal model of the scramjet established in the step (2) according to the volume dynamics principle.
For the public understanding, the technical scheme of the invention is explained in detail in the following with the accompanying drawings:
1. air inlet channel model
According to the working principle of the scramjet engine, the compression of the air inlet channel to the air flow is mainly completed through three oblique shock waves, so that the modeling of the air inlet channel of the scramjet engine is based on an oblique shock wave model.
Mach number Ma of shock wave front f And the airflow deflection angle delta can obtain a shock wave angle beta:
in the formula (I), the compound is shown in the specification,
on the basis of obtaining the shock wave angle, three oblique shock waves in the air inlet channel are sequentially solved by an oblique shock wave front-rear formula, and the outlet parameter of the air inlet channel (namely the inlet parameter of the isolation section) can be calculated by the inlet atmospheric parameter of the air inlet channel.
2. Combustion chamber model
The one-dimensional model of the scramjet engine combustion chamber regards the combustion chamber as a one-dimensional pipeline, and a one-dimensional gas dynamic equation is adopted to describe the flow condition inside the scramjet engine combustion chamber. Under the condition of given combustion chamber area change, inlet conditions and fuel adding conditions, the flow parameter distribution situation along the axial direction of the combustion chamber is obtained by solving a one-dimensional gas dynamics control equation, so that the performance parameters of the combustion chamber of the scramjet engine are estimated.
Based on experimental data, a polynomial fitting method is adopted to give total temperature distribution along the whole flow direction
Wherein x is the axial position coordinate (m) of the combustion chamber of the engine, T t (x) Is the total temperature at the x position, T t2 τ is the total heating ratio, θ is the heat release rate, and is an empirical constant from 1 to 10, χ: the non-dimensional axial position of the shaft,
x i is the location of the fuel injection point, x 4 Is the combustion chamber exit location.
Coefficient of friction prediction using the formula
In the formula, C f : local friction coefficient, k: gas adiabatic index, re x : the number of reynolds numbers in the local area,
rho is the local density, V is the local speed, x is the local axial position coordinate with the engine inlet as the origin, mu is the local aerodynamic viscosity, and the local aerodynamic viscosity is calculated by the Sutherland formula
In the formula, mu 0 Is the viscosity coefficient under standard atmospheric conditions, T s Is a Soyssen constant, related to gas properties, T c =273.16k, t is the local temperature.
Deducing to obtain a one-dimensional control equation of a flow field in a combustion chamber of the scramjet engine, namely Mach number Ma and total pressure p from a basic control equation set of gas dynamics t The variation relationship with the one-dimensional coordinate x is respectively
In the formula, A: cross-sectional area of combustion chamber, T t : total temperature; d: hydraulic diameter of the combustion chamber;
on the basis of obtaining the Mach number and the total pressure, other parameters are calculated by a gas dynamic function;
in the formula, T: static temperature, P: static pressure, V: speed, R: molar gas constant.
3. Tail nozzle model
The high-temperature and high-pressure airflow from the combustion chamber expands through the tail nozzle to generate thrust. The thermodynamic process of the gas in the nozzle can be considered as an adiabatic isentropic flow, according to the flow continuous equilibrium equation of the throat and the nozzle outlet cross section:
wherein the content of the first and second substances,
the static pressure P of the outlet of the tail nozzle can be calculated by the formula (7) 6 Velocity V 6 Flow rate of
Thus, the total thrust of the engine is obtained:
4. scramjet oblique shock mode
First, a model of the interaction of the isolation section with the combustion chamber is established. In the case of boundary layer separation in the combustion chamber, the following relationship exists between the inlet pressure and the peak pressure of the separation section of the combustion chamber
p 3 ＝δp max (34)
In the formula, p 3 Is the combustor inlet (barrier section outlet) pressure, p max To burnThe chamber peak pressure, δ, is a dimensionless similarity criterion number, and can be determined as follows: using a super-combustion shock-wave-free mode combustion chamber model, continuously increasing fuel input, enabling the lowest Mach number of the combustion chamber to be equal to 0.762 time of the Mach number of an inlet of an isolation section, namely the boundary of the super-combustion shock-wave-free mode and the super-combustion oblique shock wave mode, and calculating the ratio of the peak pressure of the combustion chamber to the outlet pressure of the isolation section at the moment;
and secondly, establishing a model of the shock wave compression section of the isolation section. On the basis of calculating the outlet pressure of the isolation section, the Mach number Ma of the outlet of the isolation section is calculated by the formula 3 Further, other parameters are obtained by the following formula:
in the formula, ma 2 To the mach number, p, of the entry of the barrier section 3 For isolating the pressure at the outlet of the section, p 2 Is the pressure at the inlet of the isolation section;
and calculating the length of the laser string in the isolated section again. Empirical relation of rectangular equal straight pipeline shock wave string length
In the formula, L s As shock string length (m), ma 2 Mach number, theta, of the starting position of the laser train 2 For the thickness (m) of the inlet boundary layer of the isolation section
Re θ Reynolds number (in terms of boundary layer momentum thickness θ) for the inlet to the isolation section 2 Characteristic length), D is the equivalent diameter (m) of the inlet of the isolation section;
then, the axial distribution p (x) of the internal pressure of the shock wave string is calculated by a cubic polynomial
Wherein χ is a dimensionless axial position,
x 2 is the location of the shock wave entry, x 3 Is the position of an outlet of the isolation section;
finally, the combustion chamber separation section model can adopt a modeling method of calculating internal parameters of a shock wave string, and the axial pressure distribution p (x) of the combustion chamber separation section is calculated by a quadratic polynomial
In the formula, χ: the non-dimensional axial position of the shaft,
x 3 is the position of the combustion chamber inlet, x max The position of the outlet of the separation section of the combustion chamber, namely the position of the highest pressure point.
5. Sub-combustion mode
The location of the critical speed of sound point is first determined. With respect to equation (7), when Ma =1, since the flow field parameters are continuous at the sonic velocity point, the following equation is certainly true regardless of the influence of the frictional force
In the formula, a represents a critical sound velocity state.
The position of the critical speed of sound point can be calculated from the above equation. The first derivative of x is simultaneously calculated for the numerator denominator to obtain:
in the formula:
solving the equation of once two-in-one can yield two solutions
According to the working principle of the scramjet engine, the airflow is accelerated from subsonic velocity to supersonic velocity in the combustion chamber, the Mach number is increased, so that the Mach number before and after the sonic velocity point can be calculated by taking a positive solution
Finally, by means of the formulae (23) and (24), ma u As an initial value condition, solving the forward integral along the axial direction of the combustion chamber from a critical sound velocity point, calculating the parameter distribution of the subsonic region of the combustion chamber by using Ma d And (5) as an initial value condition, solving by integrating backward from a critical sound velocity point along the axial direction of the combustion chamber, and calculating the parameter distribution of the supersonic velocity region of the combustion chamber. By using a modeling method of the hyper-ignition oblique shock wave modal isolation section, the parameter distribution in the isolation section can be calculated.
6. Super natural shock-free mode
When the scramjet engine is in a scramjet mode, the whole flow channel flows at supersonic speed, and the separation of the isolation section and the combustion chamber does not generate boundary layer separation. The ordinary differential equation system can be directly solved to obtain the distribution of each aerodynamic parameter of the isolation section and the combustion chamber.
7. Volume dynamic model of scramjet engine
The structure diagram of a combustion chamber of the dynamic scramjet engine built by the method is shown as 1. And (3) taking the inlet of the combustion chamber as the inlet of the cavity, taking the position of the required point as the outlet of the cavity, and carrying out volume dynamics analysis after determining the position of the outlet of the cavity. Establishing a linear ordinary differential equation of the temperature and the pressure in the cavity to obtain an update equation of the temperature and the pressure based on the time step length:
in the formula, subscript in: cavity inlet parameters, subscript out: outlet parameter of the chamber, V C : the volume of the containing cavity body is increased,
eta: combustion efficiency, W: flow rate, W f : fuel flow rate H f : fuel calorific value, h: total enthalpy, h C : the fuel static enthalpy.
The steady state result is used as an initial value, the total temperature and the total pressure are calculated by the formula (25) and the formula (26), and the Mach number is calculated by the flow formula, so that the static temperature and the static pressure are obtained. And solving by adopting a modal modeling method on the basis of obtaining the parameters, so that all the parameters of the flow field at the t moment are obtained and serve as initial values at the t +1 moment, and the calculation is carried out in a circulating manner until the total temperature and total pressure change rate is 0.
To verify the accuracy of the model, the model output was compared to experimental data. Fig. 4 and 5 are comparisons of model outputs with experimental parameters. It can be seen that the output of the model is matched with experimental parameters, the pressure peak in combustion and the pressure rising point in the isolator can be effectively captured, and the average error is lower than 3%. Through comparison of the output of the steady-state model and the experimental parameters, the established mathematical model of the scramjet engine can be proved to be used for simulating the change of the axial flow parameters in the combustion process of the scramjet engine.
Three-point fuel flow wf axially distributed in combustion chamber of engine by using mathematical model of scramjet engine as controlled object i (i =1,2, 3) is a control quantity, and the thrust force F of the engine is a controlled quantity, and a thrust instruction tracking control digital simulation is carried out to verify that the built scramjet engine can be used for designing a control system.
Under atmospheric conditions with height 27km, ma =6, a PI (proportional-integral) controller is designed, and the control effect is shown in fig. 6 and 7. As can be seen from FIG. 6, under the acceleration and deceleration control commands of large step and small step, the built scramjet engine can reflect the working state of the engine, and can be used for designing an engine control system.
The foregoing illustrates and describes the principles, general features, and advantages of the present invention. It will be understood by those skilled in the art that the present invention is not limited to the embodiments described above, which are described in the specification and illustrated only to illustrate the principle of the present invention, but that various changes and modifications may be made therein without departing from the spirit and scope of the present invention, which fall within the scope of the invention as claimed. The scope of the invention is defined by the appended claims and equivalents thereof.
Claims (4)
1. A scramjet engine volume dynamics one-dimensional modeling method is characterized by comprising the following steps: the method comprises the following steps:
step 1: establishing mathematical models of an air inlet channel, a combustion chamber and a tail nozzle of the scramjet engine according to the working principle of each part of the scramjet engine;
step 2: establishing a mathematical model of a scramjet engine in a scramjet engine scramjet mode without shock waves, a scramjet shock wave mode and a sub-combustion mode according to a scramjet engine combustion field mechanism;
and step 3: and (3) establishing a dynamic model of the scramjet based on the modal model of the scramjet established in the step (2) according to the volume dynamics principle.
2. The scramjet engine volumetric dynamics one-dimensional modeling method of claim 1, wherein: the method for establishing the mathematical model of the scramjet engine combustion chamber in the step 1 comprises the following steps:
step 1-1: based on experimental data, a polynomial fitting method is adopted to give total temperature distribution along the whole flow direction
Wherein x is the axial position coordinate (m) of the combustion chamber of the engine, T t (x) Is the total temperature at the x position, T t2 The total temperature of the inlet of the isolation section is shown, and tau is the total heating ratio; θ is the heat release rate and is an empirical constant from 1 to 10; χ: the non-dimensional axial position of the shaft,
x i is the location of the fuel injection point, x 4 Is the combustion chamber exit location;
step 1-2: coefficient of friction prediction using the formula
In the formula, C f : local friction coefficient, k: gas adiabatic index, re x : the number of reynolds numbers in the local area,
rho is the local density, V is the local velocity, x is the local axial position coordinate with the engine inlet as the origin, mu is the local aerodynamic viscosity, and the local aerodynamic viscosity is calculated by the Sutherland formula
In the formula, mu 0 Is the viscosity coefficient under standard atmospheric conditions, T s Is a Soyssinan constant, andgas properties related to, T c =273.16k, t is local temperature;
step 1-3: deducing to obtain a one-dimensional control equation of a flow field in a combustion chamber of the scramjet engine, namely Mach number Ma and total pressure p from a basic control equation set of gas dynamics t The variation relationship with the one-dimensional coordinate x is respectively
In the formula, A: cross-sectional area of combustion chamber, T t : total temperature; d: hydraulic diameter of the combustion chamber;
step 1-4: on the basis of obtaining the Mach number and the total pressure, other parameters are calculated by a gas dynamic function;
in the formula, T: static temperature, P: static pressure, V: speed, R: molar gas constant.
3. The scramjet engine volumetric dynamics one-dimensional modeling method of claim 1, wherein: the method for modeling the combustion chamber mode of the scramjet engine in the step 2 specifically comprises the following steps:
step 2-1: establishing a conversion boundary between different combustion modes, wherein the conversion boundary between the scramjet-free mode and the scramjet mode is as follows: when the lowest Mach number of the combustion chamber is larger than 0.762 times of the Mach number of the inlet of the isolation section, the combustion chamber works in a super-combustion shock-wave-free mode, and on the contrary, the combustion chamber works in a super-combustion oblique shock-wave mode; the transition boundary between the deflagration oblique shock mode and the sub-combustion mode is: when the lowest Mach number of the combustion chamber is more than or equal to 1, the combustion chamber works in a super-combustion oblique shock wave mode; on the contrary, the combustion chamber works in a sub-combustion mode;
step 2-2: when the scramjet engine is in a scramjet mode, the whole flow channel flows at supersonic speed, and the separation of the isolation section and the combustion chamber does not generate boundary layer separation; the isolation section is regarded as a one-dimensional steady fanno flow section, the combustion chamber is a variable cross-section heating pipe flow taking the geometric profile of the engine as the flow profile, and an ordinary differential equation set is solved to obtain the distribution of each pneumatic parameter of the isolation section and the combustion chamber;
step 2-3: when the scramjet engine is in a scramjet mode, the whole flow channel flows at an ultrasonic speed, and the isolation section and the combustion chamber generate boundary layer separation; the front section of the isolation section is a fannuo flow section, the rear section of the isolation section is a shock wave compression section, the front section of the combustion chamber is a combustion separation section, and the rear section of the combustion chamber is a separation attachment section; the isolation section shock wave compression section and the combustion chamber separation section take the engine geometric profile minus the separation area as a flow profile, and the isolation section Vanuno flow section and the combustion chamber separation attachment section take the engine geometric profile as the flow profile;
step 2-4: when the scramjet engine is in a sub-combustion mode, a strong shock wave string can appear in the isolation section due to high back pressure generated by the combustion chamber, the front section of the isolation section is still a fanno flow section, and the rear section of the isolation section is still a shock wave compression section; the combustion chamber inlet is subsonic, no boundary layer separation exists, and the geometric profile of the engine is used as a flow profile.
4. The scramjet engine volumetric dynamics one-dimensional modeling method of claim 1, wherein: the specific steps of establishing the dynamic model by the volume dynamics in the step 3 are as follows:
step 3-1: taking the inlet of the combustion chamber as the inlet of the cavity, taking the position of the required point as the outlet of the cavity, and carrying out volume dynamics analysis after determining the position of the outlet of the cavity;
step 3-2: establishing a linear ordinary differential equation of the temperature and the pressure in the cavity to obtain an update equation of the temperature and the pressure based on the time step length:
in the formula, the subscript in: chamber inlet parameters, subscript out: outlet parameter of the chamber, V C : the volume of the containing cavity body is increased,
eta: combustion efficiency, W: flow rate, W f : fuel flow rate H f : fuel calorific value, h: total enthalpy, h C : a static enthalpy of fuel;
step 3-3: taking the steady state result as an initial value, calculating the total temperature and the total pressure by the formulas (9) and (10), and calculating a Mach number by a flow formula to obtain a static temperature and a static pressure;
step 3-4: and (3) solving by adopting the modal modeling method in the step (2) on the basis of obtaining the parameters, so that all parameters of the flow field at the time t are obtained and serve as initial values at the time t +1, and the total pressure change rate of the total temperature is calculated in a circulating manner until the total pressure change rate of the total temperature is 0.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116562194A (en) * | 2023-07-10 | 2023-08-08 | 中国人民解放军空军工程大学 | Thrust evaluation method and system for ramjet rotary detonation engine |
| CN116562194B (en) * | 2023-07-10 | 2023-09-19 | 中国人民解放军空军工程大学 | Thrust evaluation method and system for ramjet rotary detonation engine |
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