CN109101765A - A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft - Google Patents
A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft Download PDFInfo
- Publication number
- CN109101765A CN109101765A CN201811090972.9A CN201811090972A CN109101765A CN 109101765 A CN109101765 A CN 109101765A CN 201811090972 A CN201811090972 A CN 201811090972A CN 109101765 A CN109101765 A CN 109101765A
- Authority
- CN
- China
- Prior art keywords
- mach number
- speed
- mode
- formula
- point
- 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.)
- Granted
Links
Classifications
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/10—Geometric CAD
- G06F30/15—Vehicle, aircraft or watercraft design
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Geometry (AREA)
- General Physics & Mathematics (AREA)
- Theoretical Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Computational Mathematics (AREA)
- Mathematical Analysis (AREA)
- Mathematical Optimization (AREA)
- Pure & Applied Mathematics (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- General Engineering & Computer Science (AREA)
- Automation & Control Theory (AREA)
- Testing Of Engines (AREA)
Abstract
The present invention relates to a kind of wide fast domain propulsion system modelling by mechanism methods of the big envelope curve of assembly power aircraft, the some main physical quantitys of core flow area change etc. caused by transonic flow, boundary layer are increased in existing RBCC combined dynamic system modeling method, it is close with realistic model, improves modeling accuracy.
Description
Technical field
The invention belongs to the wide fast domains of assembly power Modeling of Vehicle technical field of research more particularly to RBCC dynamical system to build
Mould technical field.The modeling method can be widely applied to all kinds of hypersonic aircraft designs and performance based on RBCC power
In analysis.
Background technique
Stream in RBCC (Rocket Based Combined Cycle) power Air-breathing hypersonic vehicle propulsion system
Road modeling technique is one of core key technology of hypersonic aircraft.RBCC engine had both remained rocket propulsion in low speed
The big thrust loading of (Rocket ejector mode, 0≤Ma≤2.5), dense atmosphere outer (pure rocket mode, 8~10≤Ma) are applicable in
Ability also has the high specific impulse and economy of punching engine.Wherein, Rocket ejector mode and pure rocket modal model can roots
It is established according to traditional conservation theory, and bimodal ultra-combustion ramjet model is as wherein important a part, is to have larger challenge
The technology of property.Dual-mode Scramjet be broadly divided into sub- combustion punching press (2.5≤Ma≤5.5) and ultra-combustion ramjet (5.5≤
Ma≤8~10) two kinds of operational modals, Modeling Research method mainly includes that theoretical mechanism modeling, numerical modeling and test are built
Mould etc..Mathematical method of the theoretical mechanism modeling as specific reflection Dual-mode Scramjet physical process, especially
It is hypersonic to air suction type to push away especially for the fast domain modeling of width of RBCC combined dynamic system due to the rapidity that it is calculated
It is of great significance into mission model development.
Limitation and higher gas velocity in RBCC Dual-mode Scramjet inner flow passage, due to geometry wall surface
Degree, on the direction perpendicular to flowing velocity, the variable gradient of flow parameter is influenced serious by flowing effects such as Three-Dimensional Shock Waves
And it is difficult to model.Therefore complexity to simplify the analysis, ignores the change perpendicular to flowing velocity direction in Fluid Control Equation
Amount forms quasi one-dimensional flow governing equation.For enable the result of quasi one-dimensional flow equation calculation be closer to it is actual as a result,
Researcher has carried out many improvement to it, including considers finite-rate reaction, fuel mixing and burning model, inner flow passage
Geometric configuration etc..
Theoretically, quasi one-dimensional flow governing equation be one-dimensional simplification to complete Fluid Control Equation as a result, and
During most of quasi- one dimensional flow analysis and research, researcher can be directed at one dimensional flow side according to its application and application range
Cheng Jinhang simplifies again, and influence of the variation of some physical quantitys (such as molecular weight, specific heat ratio) in quasi- one-dimensional stream calculation is neglected
Slightly.And in actual test data analyzer, find amount that these are simplified in RBCC Dual-mode Scramjet
It is varied widely in runner, and be directed at one dimensional flow precision of analysis and reliability to have considerable influence.Meanwhile RBCC
Transonic flow model alignment one-dimensional model integrality and the scope of application are significant in assembly power engine inner flow passage.In addition,
It is that RBCC combined dynamic system is wide that Dual-mode Scramjet model is combined with the matching of injection model and pure rocket model
The important link of fast domain modelling by mechanism.
Thus, for the fast domain characteristic of width of RBCC combined dynamic system, how to establish comprising important physical amount (such as molecule
Amount, specific heat ratio etc.), transonic flow movable model, and the Integrated Model combined with injection model and the matching of pure rocket model have
Considerable theoretical and realistic meaning.
According to the data of open source literature, current existing RBCC combined dynamic system modeling method, especially bimodal are super
The one-dimensional flow model of standard of burning ramjet is to establish on the basis of having ignored some Main physical amounts, such as transonic speed
Core flow area change caused by stream, boundary layer etc., such simplification result in the resulting result of calculating and differ with actual result
Farther out, larger impact is caused for the estimation of motor power.
Summary of the invention
Technical problems to be solved
In order to avoid the shortcomings of the prior art, the present invention proposes a kind of wide fast domain of the big envelope curve of assembly power aircraft
Propulsion system modelling by mechanism method.
Technical solution
A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft, it is characterised in that step is such as
Under:
Step 1: flight Mach number and air intake port condition for aircraft determine the mode of RBCC engine operation:
Mach number be less than 3 Ejector Mode, Mach number be greater than 3 be punching press mode, air intake port mass flow be less than air suction type propulsion
It is rocket mode when demand limits；
Step 2: for Ejector Mode, mass ratio of induced-to-inducing air n being acquired using formula (6), solves injection using formula (3)~(5)
The gas characteristic parameter c of mode outletp3、γ3、R3, the speed V of outlet is acquired using formula (8) and (11)3, static temperature T3And static pressure
p3, wherein formula (11) takes subsonic speed solution；
In formula, cpFor specific heat at constant pressure, γ is specific heat ratio, and R is gas constant,For mass flow, n is mass ratio of induced-to-inducing air, V
For flowing velocity, A is cross-sectional area, and T is temperature, and p is pressure；Subscript 1,2,3 respectively represents Rocket ejector outlet port, injection
Come outflux and mixing chamber outlet, subscript f represents secondary fuel；
Step 3: being directed to Dual-mode Scramjet model, first determine whether initial Mach number is greater than 1, if being less than
1, then mode is fired for Asia；Then be super burn mode if more than 1, using Runge-Kutta method integral ordinary differential system (14)~
(16), (17)~(23) solve unknown number ρ, V, p, Ma, T, a, F, Tt、Pt, s, while using formula (24) calculate core stream interface
Product Ac；The value of G (x) is solved using formula (26)~(27)；
In formula, ρ is density；Ma is flowing Mach number；F ' is coefficient of friction；D is hydraulic diameter；Y is speed fuel in master
Flow the component in direction and the ratio of mainstream speed；DX is the sum of following three: (1) being immersed in quiet in fluid in control volume boundary
The only resistance of object, (2) drop and liquid mist are because movement velocity is lower than the resistance that mainstream speed generates, (3) gravity or other power
Act on the power opposite with directional velocity generated in control volume；DQ is reaction heat release；dWxThe sum of conduct heat and do work for wall surface；
DH is enthalpy variation；F is axial force；TtFor total temperature；PtFor stagnation pressure；S is specific entropy；For universal gas constant；Component i's
Molar fraction Xi=Ni/ N, N are molal quantity；N is component of mixture number；
Judge the relationship of G (x) Yu Ma, and then obtains dMa2The changing rule of/dx；Specifically, dMa2/ dx with G (x) and
The changing rule of Ma is as shown in the table:
So that Ma=1 and the reference of the initial subsonic speed Mach 2 ship of G (x)=0, when initial subsonic speed Mach number is higher than this
When value, Mach number can reach Sonic Point before G (x) is equal to 0, be jammed in this point generation；When Mach number is less than this value,
Sonic Point has not yet been reached in Mach number when G (x) is equal to 0, and Mach number can reach its subsonic speed maximum value in G (x)=0 point, along the side x
Downstream Mach number is gradually reduced；
The present invention only considers that Mach number can be by the state of Sonic Point, by taking subsonic speed to supersonic transition as an example, when wanting
When Mach number being asked smoothly to accelerate to supersonic speed, dMa2/ dx must be positive in subsonic speed and supersonic speed Duan Jun, i.e., Mach number gradually increases
Add；As seen from the above table, G (x) is necessarily less than zero in subsonic speed region, and having to be larger than zero in supersonic speed region just can guarantee Mach number
Increase；But when Mach number is equal to 1, non trivial solution is in a saddle point mathematically, and the different values of G (x) correspond to
Three kinds of different results:
1.G(x)<0；Mach number will approach 1 and no longer will increase or reduce；
2.G(x)>0；Mach number forever can not be close to 1；
3.G (x)=0；Mach number saddle point can be passed through using jump method；In fact, the value of G (x) is rigid as Ma=1
Benefit is at zero point, this point, dMa2/ dx is uncertain, thus after this point, Mach number is possible to accelerate to ultrasound
Speed, it is also possible to which subsonic speed of slowing down back, test think that supersonic speed direction is taken more to tally with the actual situation；
For first two state, need to adjust initial parameter to recalculate, until meeting state 3 or condition of being jammed,
To acquire engine export aerodynamic parameter ρ, V, p, Ma, T, a, F, T of punching press modet、Pt, s value；
Step 4: thrust and ratio for rocket mode, when calculating separately rocket mode according to formula (32) and formula (36)
Punching obtains the performance indicator of rocket mode:
Beneficial effect
A kind of wide fast domain propulsion system modelling by mechanism method of the big envelope curve of assembly power aircraft proposed by the present invention, with
Test (William H.Heiser, David T.Pratt, Daniel H.Daley, the et al.Hypersonic of Billig
Airbreathing propulsion [M] .Washington, D.C.:AIAA, 1994, pp.365-367) it is reference, in conjunction with height
Precision CFD emulation, verifies the precision of established model, as shown in Figure 4.It can be seen that the model established is in exhausted big portion
There is good approximation ratio to test result and CFD simulation result within the scope of point.And since one-dimensional model cannot be handled
Situations such as reflection of Three-Dimensional Shock Wave string, intersection, thus following closely in a distance after propellant spray, the model calculation cannot be quasi-
True tracking test measured value, even so, the model established in entire runner, pressure distribution and test measurements
Mean square error is only 1.67%.
Detailed description of the invention
Fig. 1 injection engine configuration schematic diagram
Fig. 2 RBCC power aerial vehicle propulsion system Longitudinal cross section schematic
Fig. 3 rocket engine schematic diagram
Fig. 4 model verification result figure
Flow chart Fig. 5 of the invention
Specific embodiment
Now in conjunction with embodiment, attached drawing, the invention will be further described:
The purpose of the present invention is being directed to the modeling of RBCC combined dynamic system, a kind of consideration Ejector Mode, bimodal are proposed
The fast domain propulsion system model of the width of scramjet engine mode and rocket mode.
The purpose of the present invention is realized by following modeling method:
Ejector Mode model and pure rocket modal model are established using the Ejector Mode modeling method based on law of conservation,
It is characterized in that: can reflect the physical characteristic of Ejector Mode and calculating speed is very fast.Using consideration specific heat ratio, boundary layer and across sound
The quasi- one dimensional flow of fast flow characteristics is theoretical, establishes wide fast domain Dual-mode Scramjet model, can guarantee certain essence
Physics law on the basis of degree, in reflection inner flow passage flowing.
Assembly power model proposed by the invention is suitable for the hypersonic propulsion system of air suction type based on RBCC.
Novel compositions dynamic model overall procedure proposed by the invention are as follows:
(1) Ejector Mode model.For injection engine configuration as shown in Figure 1, it is assumed that all gas is uniform
Calorimetric ideal gas, and complex three-dimensional effect is not considered.
It is defined as follows parameter first:
In formula, RareaFor injection air-flow and by the ratio between injection flow cross-sectional product；Supporting plate blockage ratio is represented, A is cross section
Product, subscript 1,2,3 respectively represents Rocket ejector outlet port, injection carrys out outflux and mixing chamber outlet (similarly hereinafter).
Due to injection/mixing chamber gas components unchanged, that is, following mixing principle can be used:
In formula, cpFor specific heat at constant pressure, γ is specific heat ratio, and R is characterized gas constant,For mass flow, n is injection system
Number (is only limitted to Ejector Mode mark):
According to injector outlet and inlet mass conservation:
Subscript f represents secondary fuel in formula, and V is speed, and T is temperature, and p is pressure.By front it is assumed that the conservation of momentum
Are as follows:
In formula, subscript " ' " refer to that speed fuel or area of injection orifice in the projection in mainstream speed direction, ignore wall friction τw
When with secondary fuel addition:
Corresponding energy conservation equation are as follows:
In formula, TtFor total temperature；W be major flow cross wall surface externally do work, outer bound pair mainstream acting and spray fuel and
Heat transfer caused by the temperature difference of mainstream；Q is the burning heat release of fuel and oxide in mainstream.Ignore acting and secondary fuel burning
When:
Given injector entrance once flows and the parameter p of Secondary Flow1、T1、Tt1、R1、γ1、cp1、A1、V1And p2、 T2、
Tt2、R2、γ2、cp2、A22、V2, mass ratio of induced-to-inducing air n is acquired according to formula (6), is entered according to formula (3)~(5) available injector
The gas characteristic parameter c of mouthp3、γ3、R3.Can acquire the total temperature of outlet by formula (13), so by total temperature, static temperature and speed it
Between relation equation (basic Euler equations), be brought into solved in formula (8) and (11) injector outlet speed V3, static temperature
T3With static pressure p3, that is, obtain whole flow parameters of injector outlet.It should be noted that formula (11) are one about speed
First quadratic equation, the speed solved there are two value, one be subsonic speed, another be supersonic speed, and before the two is normal shock wave
Relationship afterwards, entropy corresponding to subsonic speed solution is higher, thus it is big to be considered a possibility that taking subsonic speed solution.
(2) Dual-mode Scramjet model.It is vertical to RBCC power aerial vehicle propulsion system as shown in Figure 2 to cut
Face, the engine portion in dotted line frame is divided into inner flow passage of interest, including straight sections such as isolator, bimodal combustion chambers
With linear expansion section.Assuming that flowing has the feature that rocket serves as the effect of propellant spray device in punching press mode, ignore rocket
The influence of the internal flow channel cross-section product of supporting plate；Flowing is quasi- one-dimensional and is the steady motion of a fluid；The variation of flow parameter is continuous；Gas
Stream belongs to semi ideal air-flow, that is, follows Boyle and Charles criterion, and specific heat ratio is only the function of temperature and gas component.
Taking a certain infinitesimal length along flowing velocity direction is the flowing control volume of dx, for this control volume, continuity equation,
The equation of momentum and energy equation are expressed as following differential form:
In formula:For mass flow；ρ is density；A is cross-sectional area；V is flowing velocity；P is pressure；γ is specific heat ratio；
Ma is Mach number；F is coefficient of friction；D is hydraulic diameter；Y is speed fuel in the component of main flow direction and the ratio of mainstream speed
Value；DX is the sum of following three: (1) being immersed in the resistance of the stationary object in control volume boundary in fluid, (2) drop and liquid mist
Because movement velocity is lower than the resistance that mainstream speed generates, (3) gravity or other power act on generate in control volume and speed
Contrary power.In energy equation (16): dQ is reaction heat release； dWxThe sum of conduct heat and do work for wall surface；DH is enthalpy change
Change；cpFor specific heat at constant pressure；T is temperature.
The equation of gas state are as follows:
In formula: W is molecular mass (air 28.92).According to the relationship of the velocity of sound and airflow characteristic parameter, velocity of sound a's is micro-
Divide expression formula are as follows:
For Mach number, axial force, total temperature, stagnation pressure, specific entropy, there is the expression of following differential form:
In formula: F is axial force；TtFor total temperature；PtFor stagnation pressure；S is specific entropy；R is characterized gas constant；For argoshield
Constant, 8.314472J/ (Kmol)；The molar fraction X of component ii=Ni/ N (N is molal quantity)；N is mixing
Object number of components (is only limitted to Dual-mode Scramjet model).10 differential such as formula (14)~(16), (17)~(23)
Equation shares 10 unknown numbers ρ, V, p, Ma, T, a, F, Tt、Pt, s, can be solved by Runge-Kutta method.
Influence of the boundary layer to flowing is very big in RBCC Dual-mode Scramjet inner flow passage, is that engine mode turns
One of principal element changed, since boundary layer belongs to low speed Low Energy Region, it is assumed herein that momentum and speed all in runner by
Core flow undertakes, and identical pressure and temperature is shared in boundary layer with mainstream.Take AcFor core flow sectional area, by formula (15),
(17) it brings formula (14) into (21) and the differential equation of the core flow sectional area about flow parameter can be obtained by arrangement:
For the transonic flow in RBCC engine inner flow passage, convolution (14)~(16), (17), (19) are available
Mach number evolution with distance are as follows:
It enables:
Then:
Wherein:
(3) pure rocket modal model.Rocket engine is illustrated as shown in figure 3, fuel and oxidant are respectively via respective
Pipeline enters combustion chamber, and violent and rapid chemical combustion reaction, in the most areas of combustion chamber, combustion occurs in combustion chamber
The flow velocity of product is burnt all close to zero, thus its temperature and pressure is close to total temperature and stagnation pressure.Combustion product is shunk by throat
Section ramps up, and until reaching local velocity of sound in throat, then enters jet pipe and further expands acceleration, until nozzle velocity
Ve。
For the perfect gas of constant specific heat capacity, by the available specific heat at constant pressure c of the definition of enthalpypWith specific heat ratio γ and spy
Levy the relationship of gas constant R are as follows:
For perfect gas, by energy equation, total temperature TtHave between temperature T:
In formula, speed V=Maa, velocity of sound a2=γ RT, and stagnation pressure PtIn the relationship of pressure p are as follows:
In addition, the thrust P of engine is equal to inner surface pressure p according to the conservation of momentum and Newton's second lawsIn propulsion side
To component and outer surface pressure p∞In the sum of direction of propulsion component:
In formula, S is area element,For the sum of oxidant and fuel mass flow rates, Ve、AeRespectively spray
The air velocity and cross-sectional area of pipe outlet.
According to combustion chamber and the nozzle exit conservation of energy, have:
In formula, Tt,cFor total temperature in combustion chamber.Solve Ve:
There is relational expression p between temperature and pressure0/p1=(T0/T1)γ/(γ-1), and equation (29) are substituted into equation (34):
The propulsive efficiency of engine can indicate by specific impulse, and specific impulse is defined as pushing away for per unit mass stream under sea-level condition
Power, if it is assumed that the pressure of nozzle exit section is equal to environmental pressure, i.e. jet pipe reaches fully expanded state, then equation (32)
Middle pe=p∞, then:
Below with reference to flow chart, embodiments of the present invention will be described.
Step 1: the flight Mach number and air intake port condition for aircraft determine the mould of RBCC engine operation
State: Mach number is that be greater than 3 be that punching press mode, air intake port mass flow are pushed away less than air suction type for Ejector Mode, Mach number less than 3
It is rocket mode when being limited into demand；
Step 2: acquiring mass ratio of induced-to-inducing air n using formula (6) for Ejector Mode, injection is solved using formula (3)~(5)
The gas characteristic parameter c of mode outletp3、γ3、R3, can acquire the total temperature of outlet by formula (13), so by total temperature, static temperature and
Relation equation (basic Euler equations) between speed, the speed V of outlet is acquired in conjunction with formula (8) and (11)3, static temperature T3With it is quiet
Press p3(wherein formula (11) takes subsonic speed solution)；
Step 3: being directed to Dual-mode Scramjet model, first determine whether initial Mach number is greater than 1, if small
In 1, then mode is fired for Asia.Then it is super burn mode if more than 1, ordinary differential system (14) is integrated using Runge-Kutta method
~(16), (17)~(23) (core flow area using formula (24) calculate, while in integral process cross-sectional area A with core
Flow section accumulates AcSubstitution, in terms of and boundary layer influence), solve unknown number ρ, V, p, Ma, T, a, F, Tt、Pt,s.In addition entire
The value of formula (26)~(27) estimate simultaneously G (x) is used in integral process, and judges its relationship with Ma, and then obtain dMa2/
The changing rule of dx.Specifically, dMa2/ dx is as shown in the table with the changing rule of G (x) and Ma:
As an example it is supposed that certain initial subsonic speed Mach number makes G (x)=0 when being integrated to Ma=1 just, when initial
When subsonic speed Mach number is higher than this value, Mach number can reach Sonic Point before G (x) is equal to 0, be jammed in this point generation；When
When Mach number is less than this value, when G (x) is equal to 0, Sonic Point is had not yet been reached in Mach number, and Mach number can reach it in G (x)=0 point
Subsonic speed maximum value, downstream Mach number is gradually reduced in the x-direction.
In the present invention only consider Mach number can by the state of Sonic Point, by taking subsonic speed to supersonic transition as an example, when
It is required that when Mach number smoothly accelerates to supersonic speed, dMa2/ dx must be positive in subsonic speed and supersonic speed Duan Jun, i.e., Mach number is gradually
Increase.As seen from the above table, G (x) is necessarily less than zero in subsonic speed region, and having to be larger than zero in supersonic speed region just can guarantee Mach
Several increases.But when Mach number is equal to 1, non trivial solution is in a saddle point mathematically, and the different values of G (x) are corresponding
In three kinds of different results:
1.G(x)<0.Mach number will approach 1 and no longer will increase or reduce；
2.G(x)>0.Mach number forever can not be close to 1；
3.G (x)=0.Mach number saddle point can be passed through using jump method.In fact, the value of G (x) is rigid as Ma=1
Benefit is at zero point, this point, dMa2/ dx is uncertain, thus after this point, Mach number is possible to accelerate to ultrasound
Speed, it is also possible to which subsonic speed of slowing down back, test think that supersonic speed direction is taken more to tally with the actual situation.
For first two state, need to adjust initial parameter to recalculate and (restart third step), until meeting shape
State 3 or condition of being jammed, to acquire engine export aerodynamic parameter ρ, V, p, Ma, T, a, F, T of punching press modet、Pt, s
Value；
Step 4: for rocket mode, thrust P when calculating separately rocket mode according to formula (32) and formula (36) and
Specific impulse Isp, obtain the performance indicator of rocket mode motor power and specific impulse.
Step 5: the configuration for given engine inner flow passage designs, engine is solved in difference using model above
The performance parameters such as thrust and specific impulse under flying condition, according to regular (the multimode engine of preset mode switched control
Generally can be used can only use overall conversion by module conversion or overall conversion, single module engine), to assess stream in engine
Big fast domain characteristic in the entire flight envelope in road, conclusion is for instructing engine design optimization and control.
Claims (1)
1. a kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft, it is characterised in that step is such as
Under:
Step 1: flight Mach number and air intake port condition for aircraft determine the mode of RBCC engine operation: Mach
Number be less than 3 Ejector Mode, Mach number be greater than 3 be punching press mode, air intake port mass flow be less than air suction type promote demand
It is rocket mode when limitation；
Step 2: for Ejector Mode, mass ratio of induced-to-inducing air n being acquired using formula (6), solves Ejector Mode using formula (3)~(5)
The gas characteristic parameter c of outletp3、γ3、R3, the speed V of outlet is acquired using formula (8) and (11)3, static temperature T3With static pressure p3,
Wherein formula (11) takes subsonic speed solution；
In formula, cpFor specific heat at constant pressure, γ is specific heat ratio, and R is gas constant,For mass flow, n is mass ratio of induced-to-inducing air, and V is stream
Dynamic speed, A is cross-sectional area, and T is temperature, and p is pressure；Subscript 1,2,3 respectively represents Rocket ejector outlet port, injection incoming flow
Outlet and mixing chamber outlet, subscript f represent secondary fuel；
Step 3: it is directed to Dual-mode Scramjet model, first determines whether initial Mach number is greater than 1, if less than 1,
Mode is fired for Asia；Then be super burn mode if more than 1, using Runge-Kutta method integral ordinary differential system (14)~
(16), (17)~(23) solve unknown number ρ, V, p, Ma, T, a, F, Tt、Pt, s, while using formula (24) calculate core stream interface
Product Ac；The value of G (x) is solved using formula (26)~(27)；
In formula, ρ is density；Ma is flowing Mach number；F ' is coefficient of friction；D is hydraulic diameter；Y is speed fuel in mainstream side
To component and mainstream speed ratio；DX is the sum of following three: (1) being immersed in the resting in control volume boundary in fluid
The resistance of body, (2) drop and liquid mist are because movement velocity is lower than the resistance that mainstream speed generates, (3) gravity or the effect of other power
The power opposite with directional velocity generated in control volume；DQ is reaction heat release；dWxThe sum of conduct heat and do work for wall surface；DH is
Enthalpy variation；F is axial force；TtFor total temperature；PtFor stagnation pressure；S is specific entropy；For universal gas constant；Mole of component i
Score Xi=Ni/ N, N are molal quantity；N is component of mixture number；
Judge the relationship of G (x) Yu Ma, and then obtains dMa2The changing rule of/dx；Specifically, dMa2/ dx is with G's (x) and Ma
Changing rule is as shown in the table:
So that Ma=1 and the reference of the initial subsonic speed Mach 2 ship of G (x)=0, when initial subsonic speed Mach number is higher than this value
When, Mach number can reach Sonic Point before G (x) is equal to 0, be jammed in this point generation；When Mach number is less than this value, in G
(x) Sonic Point has not yet been reached in Mach number when being equal to 0, and Mach number can reach its subsonic speed maximum value in G (x)=0 point, along the side x
Downstream Mach number is gradually reduced；
The present invention only considers that Mach number can be by the state of Sonic Point, by taking subsonic speed to supersonic transition as an example, when requiring horse
When conspicuous number smoothly accelerates to supersonic speed, dMa2/ dx must be positive in subsonic speed and supersonic speed Duan Jun, i.e., Mach number gradually increases；
As seen from the above table, G (x) is necessarily less than zero in subsonic speed region, and having to be larger than zero in supersonic speed region just can guarantee Mach number
Increase；But when Mach number is equal to 1, non trivial solution is in a saddle point mathematically, and the different values of G (x) correspond to three
The different result of kind:
1.G(x)<0；Mach number will approach 1 and no longer will increase or reduce；
2.G(x)>0；Mach number forever can not be close to 1；
3.G (x)=0；Mach number saddle point can be passed through using jump method；In fact, the value of G (x) is just located as Ma=1
At zero point, this point, dMa2/ dx is uncertain, thus after this point, Mach number is possible to accelerate to supersonic speed,
It is possible that subsonic speed of slowing down back, test thinks that supersonic speed direction is taken more to tally with the actual situation；
For first two state, need to adjust initial parameter to recalculate, until meeting state 3 or condition of being jammed, thus
Acquire engine export aerodynamic parameter ρ, V, p, Ma, T, a, F, T of punching press modet、Pt, s value；
Step 4: for rocket mode, thrust and specific impulse when rocket mode are calculated separately according to formula (32) and formula (36),
Obtain the performance indicator of rocket mode:
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811090972.9A CN109101765B (en) | 2018-09-19 | 2018-09-19 | Mechanism modeling method for large-package-width fast-domain propulsion system of combined power aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811090972.9A CN109101765B (en) | 2018-09-19 | 2018-09-19 | Mechanism modeling method for large-package-width fast-domain propulsion system of combined power aircraft |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109101765A true CN109101765A (en) | 2018-12-28 |
| CN109101765B CN109101765B (en) | 2022-06-14 |
Family
ID=64866695
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201811090972.9A Active CN109101765B (en) | 2018-09-19 | 2018-09-19 | Mechanism modeling method for large-package-width fast-domain propulsion system of combined power aircraft |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109101765B (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110362783A (en) * | 2019-05-31 | 2019-10-22 | 中国人民解放军国防科技大学 | Rocket-based combined cycle engine thrust calculation method |
| CN111271326A (en) * | 2020-01-16 | 2020-06-12 | 集美大学 | Supersonic ejector design and evaluation method |
| CN114492227A (en) * | 2021-12-29 | 2022-05-13 | 中国航天空气动力技术研究院 | Method, device and medium for analyzing dynamic stability based on aircraft oscillation frequency |
| CN114722743A (en) * | 2022-05-24 | 2022-07-08 | 中国人民解放军国防科技大学 | Combustion chamber chemical balance-based scramjet engine one-dimensional performance estimation method |
| CN114912202A (en) * | 2022-05-24 | 2022-08-16 | 大连理工大学 | Integrated coupling control method for propelling of wide-speed-range air-breathing power aircraft body |
| CN116341407A (en) * | 2023-03-10 | 2023-06-27 | 城林科技(上海)有限公司 | Aerodynamic design method for test bed |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8256203B1 (en) * | 2007-01-26 | 2012-09-04 | The University Of Alabama In Huntsville | Rocket based combined cycle propulsion unit having external rocket thrusters |
| CN103758663A (en) * | 2014-01-20 | 2014-04-30 | 西北工业大学 | Rocket-based combined cycle ejector mode performance test engine |
| CN106050472A (en) * | 2016-07-08 | 2016-10-26 | 西北工业大学 | Turbo-rocket combined ramjet engine and operating method thereof |
| CN106150757A (en) * | 2016-08-10 | 2016-11-23 | 西北工业大学 | A kind of dual pathways becomes geometry rocket based combined cycle electromotor |
| CN107061010A (en) * | 2017-03-23 | 2017-08-18 | 西北工业大学 | A kind of rocket based combined cycle engine structure changes air intake duct |
| CN207093230U (en) * | 2017-06-13 | 2018-03-13 | 南昌航空大学 | A kind of monoblock type rocket and ultra-combustion ramjet combined engine |
-
2018
- 2018-09-19 CN CN201811090972.9A patent/CN109101765B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8256203B1 (en) * | 2007-01-26 | 2012-09-04 | The University Of Alabama In Huntsville | Rocket based combined cycle propulsion unit having external rocket thrusters |
| CN103758663A (en) * | 2014-01-20 | 2014-04-30 | 西北工业大学 | Rocket-based combined cycle ejector mode performance test engine |
| CN106050472A (en) * | 2016-07-08 | 2016-10-26 | 西北工业大学 | Turbo-rocket combined ramjet engine and operating method thereof |
| CN106150757A (en) * | 2016-08-10 | 2016-11-23 | 西北工业大学 | A kind of dual pathways becomes geometry rocket based combined cycle electromotor |
| CN107061010A (en) * | 2017-03-23 | 2017-08-18 | 西北工业大学 | A kind of rocket based combined cycle engine structure changes air intake duct |
| CN207093230U (en) * | 2017-06-13 | 2018-03-13 | 南昌航空大学 | A kind of monoblock type rocket and ultra-combustion ramjet combined engine |
Non-Patent Citations (2)
| Title |
|---|
| ANKIT GOEL等: "Retrospective cost adaptive thrust control of a 1D scramjet with Mach number disturbance", 《2015 AMERICAN CONTROL CONFERENCE (ACC)》 * |
| 吕翔 等: "RBCC飞行器爬升段轨迹设计方法", 《航空学报》 * |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110362783A (en) * | 2019-05-31 | 2019-10-22 | 中国人民解放军国防科技大学 | Rocket-based combined cycle engine thrust calculation method |
| CN111271326A (en) * | 2020-01-16 | 2020-06-12 | 集美大学 | Supersonic ejector design and evaluation method |
| CN114492227A (en) * | 2021-12-29 | 2022-05-13 | 中国航天空气动力技术研究院 | Method, device and medium for analyzing dynamic stability based on aircraft oscillation frequency |
| CN114722743A (en) * | 2022-05-24 | 2022-07-08 | 中国人民解放军国防科技大学 | Combustion chamber chemical balance-based scramjet engine one-dimensional performance estimation method |
| CN114912202A (en) * | 2022-05-24 | 2022-08-16 | 大连理工大学 | Integrated coupling control method for propelling of wide-speed-range air-breathing power aircraft body |
| CN116341407A (en) * | 2023-03-10 | 2023-06-27 | 城林科技(上海)有限公司 | Aerodynamic design method for test bed |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109101765B (en) | 2022-06-14 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN109101765A (en) | A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft | |
| CN111339681B (en) | Nozzle outlet parameter matching method for simulating engine gas medium jet flow aerodynamic interference effect by adopting air medium | |
| CN101539480B (en) | One-dimensional evaluation method of combustion efficiency for scramjet engine | |
| Slater et al. | Modeling of fixed-exit porous bleed systems for supersonic inlets | |
| Shi et al. | Numerical analysis of flow features and operation characteristics of a rocket-based combined-cycle inlet in ejector mode | |
| CN115688287A (en) | Design method of aviation turbofan engine with ejector nozzle | |
| Gamble et al. | Improving off-design nozzle performance using fluidic injection | |
| Zhang et al. | Experimental and numerical investigation of a fluidically variable hypersonic inlet | |
| Rao et al. | Effect of nozzle geometry on flowfield for high subsonic jets | |
| Khalid et al. | Computational fluid dynamics analysis of compressible flow through a converging-diverging nozzle using the k-ε turbulence model | |
| Wu et al. | A fluidic thrust vector control using the bypass flow in a dual throat nozzle | |
| Etele et al. | Experiments with ejector rocket entrainment | |
| Wise et al. | Experimental investigation of a three-dimensional scramjet engine at Mach 12 | |
| Kanda et al. | Experimental study of a combined-cycle engine combustor in ejector-jet mode | |
| Lin et al. | Effects of fuel-lean primary rocket on bypass ratio in RBCC ejector mode | |
| Yu et al. | Numerical Investigation on Mechanism of External Flow Field–Induced Separation Pattern Transition | |
| Erdem et al. | Parametric study of secondary gas injection into a conical rocket nozzle for thrust vectoring | |
| El-Zahaby et al. | Study of the configuration and performance of air-air ejectors based on cfd simulation | |
| Kumar et al. | Nozzle design of oblique detonation wave engine | |
| Bengherbia et al. | Numerical investigation of thermally choked ram accelerator in sub-detonative regime | |
| Sivakumar et al. | Numerical simulations of flow in a 3-D supersonic intake at high Mach numbers | |
| Thangadpai et al. | Numerical investigation of the intake flow characteristics for a ramjet engine with and without heat addition in the combustion chamber | |
| Bian et al. | Numerical Simulations of Two-Dimensional Hypersonic Inlet At Finite Flight Angles | |
| Gronland et al. | Nozzle/afterbody integration of hypersonic vehicles by means of secondary air injection | |
| Noh | Numerical Investigation of Chocked Converging-Diverging Nozzles for Thruster Application |
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 |