CN115688287A - Design method of aviation turbofan engine with ejector nozzle - Google Patents
Design method of aviation turbofan engine with ejector nozzle Download PDFInfo
- Publication number
- CN115688287A CN115688287A CN202211432776.1A CN202211432776A CN115688287A CN 115688287 A CN115688287 A CN 115688287A CN 202211432776 A CN202211432776 A CN 202211432776A CN 115688287 A CN115688287 A CN 115688287A
- Authority
- CN
- China
- Prior art keywords
- flow
- ratio
- engine
- nozzle
- injection
- 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.)
- Pending
Links
Images
Classifications
-
- 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
- Testing Of Engines (AREA)
Abstract
The invention discloses a design method of an aviation turbofan engine with an injection nozzle. The invention improves a parameter cycle analysis model, and in the improved parameter cycle analysis model, the internal airflow flow of the injection nozzle in a design state meets the following requirements: (1) The main flow and the secondary flow are accelerated to an outlet supersonic speed state from an inlet subsonic speed state respectively, and meanwhile, the two flows of gas are in a complete expansion state at the outlet of the spray pipe; (2) the two gases do not mix in the spray pipe and flow in an isentropic manner; and (3) keeping static pressure balance at the junction of the primary flow and the secondary flow all the time. Compared with the prior art, the invention provides a multistage iteration method which uses a bypass ratio, an injection bypass ratio, a fan pressure ratio, a compressor pressure ratio and a combustion chamber outlet temperature as iteration variables and uses a bypass airflow state, an injection secondary flow state, engine thrust and oil consumption rate as inspection indexes, and the aviation turbofan engine with good matching relationship between an injection nozzle and a core engine and excellent overall performance can be designed based on the final iteration result.
Description
Technical Field
The invention relates to a design method of an aviation turbofan engine, in particular to a design method of an aviation turbofan engine with an injection nozzle.
Background
The small bypass ratio turbofan engine is used as a main power source of a modern military warplane, and an exhaust system of the small bypass ratio turbofan engine has a crucial influence on the thrust and infrared characteristics of the engine. The ejector nozzle is widely applied to the field of airplane design due to the advantages of simple structure, small thrust loss, capability of effectively reducing the infrared radiation intensity and the like. 20 fighters in China, B-2 bombers in America and YF-23 fighters in China all adopt an injection exhaust technology.
The design method of the gas ejector mainly comprises a one-dimensional flow method, a characteristic line method and a CFD numerical simulation method. In 1939, fluguel firstly provides a one-dimensional analysis method of the ejector by applying a continuous equation, a momentum conservation equation and an energy conservation equation, and then Keenan and Neumann simplify and perfect the one-dimensional analysis method, thereby laying the foundation of the one-dimensional design theory of the ejector. However, the one-dimensional analysis theory of the ejector introduces a plurality of simplifying assumptions, and the accuracy and the applicability of the theory limit the development of the theory. A characteristic line method is a main design method of an early supersonic velocity spray pipe, a classical characteristic line grid method is proposed by Busemann in 1931, intersection and reflection of expansion waves inside the spray pipe are calculated by utilizing a characteristic line theory, and a smooth spray pipe molded line is finally formed. The latter researchers developed various analytical methods of nozzle design based on the characteristic line theory, including the Foelsch method, the Cresci method, the Silvells method, and the like. The methods fill the defects of the classical characteristic line method, are widely applied to engineering, but have difficulty in obtaining better results for complex flows such as shock wave intersection, turbulent flow mixing and the like. With the development of computer science and technology, a CFD numerical simulation method can well simulate complex flow in a flow field, at present, a learner prefers to adopt commercial numerical simulation software to design the profile of the injection nozzle in an expansion mode, but a fine numerical simulation process usually takes a lot of time.
The domestic design research on the ejector nozzle mainly focuses on exploring the influence of different design parameters on the flow characteristics, thrust, infrared and other performances of the nozzle. The Liufu city and the like adopt a numerical simulation method to research and obtain the influence rule of the distance ratio change and the area ratio change of the binary injection spray pipes on the thrust characteristic and the infrared characteristic; the Dengdonjian and the like design various injection nozzle schemes by changing the diameter and the length of an outlet of an outer sleeve and explore the flow and thrust performance of the Dengdonjian in various states by adopting CFX; regular researches are carried out aiming at main profile design parameters of the injection nozzle at design points by Zhang Ke et al, and the influence of different profile parameters on the characteristics and the thrust performance of the internal flow shear layer is explored.
The above studies are limited to the design and simulation of the nozzle as a separate component, and there are no clear bases and limitations for the selection of the structural dimensions and boundary conditions of the nozzle, which results in test results that are not representative of the actual performance of the nozzle mounted on the engine. In engineering practice, the problem that the designed nozzle is poor in matching with an engine core engine generally exists, and the specific characteristics are that the nozzle cannot provide expected thrust after being installed and can have adverse effects on the thermodynamic cycle of the engine core engine. Therefore, the research on the matching relation between the injection nozzle and the core engine in the overall design stage of the engine has important significance on the design and test of the nozzle, and the research and development cost can be effectively reduced.
Disclosure of Invention
The invention aims to overcome the technical problems in the prior art and provide a design method of an aviation turbofan engine with an ejector nozzle, wherein the design method is characterized in that a parameter cycle analysis method is improved, and the flow matching between the ejector nozzle and a core engine and the thrust and oil consumption performance of the engine are comprehensively considered in the design process, so that the aviation turbofan engine with the core engine and the ejector nozzle having a good matching relationship and excellent overall performance can be designed.
The invention specifically adopts the following technical scheme to solve the technical problems:
a design method of an aviation turbofan engine with an ejector nozzle comprises the steps of obtaining a two-dimensional geometric structure of the ejector nozzle based on an improved parameter cycle analysis model; in the improved parameter cycle analysis model, the airflow flow inside the injection nozzle meets the following conditions of the Mach number of the main flow and the secondary flow and the static pressure distribution:
where M represents the section Mach number, P represents the section static pressure, the subscripts 7, 8, 9 represent the main flow inlet, throat and outlet sections, the subscripts 27, 28, 29 represent the secondary flow inlet, throat and outlet sections, P represents the section static pressure, and b representing ambient backpressure.
Preferably, the model input variables of the improved parameter cycle analysis model comprise an injection bypass ratio and a ratio of static pressures of the primary outlet and the secondary outlet to environmental backpressure.
Preferably, in the process of selecting the design value of the engine core component, a multi-stage iteration method which takes a bypass ratio, an injection bypass ratio, a fan pressure ratio, a compressor pressure ratio and the outlet temperature of a combustion chamber as iteration variables and takes a bypass airflow state, an injection secondary flow state, engine thrust and oil consumption rate as inspection indexes is used.
Compared with the prior art, the invention has the following beneficial effects:
the improved parameter cycle analysis model is introduced in the design process of the aviation turbofan engine with the ejector nozzle, and the flow matching between the ejector nozzle and the core engine and the thrust and oil consumption performance of the engine are comprehensively considered, so that the aviation turbofan engine with the core engine and the ejector nozzle having a good matching relationship and excellent thrust performance can be designed.
Drawings
FIG. 1 is a structural diagram of a turbofan engine with fan bypass injection
FIG. 2 is a flow chart of the calculation of the jet nozzle block
FIG. 3 is a flow chart of iterative selection of engine core components
FIG. 4 is a structural view of a designed ejector nozzle;
FIG. 5 is a drawing of a jet nozzle and outer flow field grid division;
FIG. 6 is a schematic view of a flow field calculation domain;
FIG. 7 is a graph of the change in the draw coefficient and thrust coefficient of a nozzle at a plurality of typical mission points;
FIG. 8 is a graph showing the variation of the thrust coefficient with the pressure drop ratio of the nozzle under the condition of constant injection coefficient.
Detailed Description
Aiming at the defects existing in the prior art that the ejector nozzle is designed as an independent part, the solution idea of the invention is to improve the traditional parameter cycle analysis method and comprehensively consider the flow matching between the ejector nozzle and the core engine and the thrust and oil consumption performance of the engine in the design process, so that the aviation turbofan engine with good matching relationship between the core engine and the ejector nozzle and excellent thrust performance can be designed. .
The technical scheme provided by the invention is as follows:
a design method of an aviation turbofan engine with an ejector nozzle comprises the steps of obtaining a two-dimensional geometric structure of the ejector nozzle based on an improved parameter cycle analysis model; in the improved parameter cycle analysis model, the airflow flow inside the injection nozzle meets the following conditions of the Mach number of the main flow and the secondary flow and the static pressure distribution:
where M represents the section Mach number, P represents the section static pressure, the subscripts 7, 8, 9 represent the main flow inlet, throat and outlet sections, the subscripts 27, 28, 29 represent the secondary flow inlet, throat and outlet sections, P represents the section static pressure, and b representing ambient backpressure.
Preferably, the model input variables of the improved parameter cycle analysis model comprise an injection bypass ratio and a ratio of static pressures of the primary outlet and the secondary outlet to environmental backpressure.
Preferably, in the process of selecting the design value of the engine core component, a multi-stage iteration method which takes a bypass ratio, an injection bypass ratio, a fan pressure ratio, a compressor pressure ratio and the outlet temperature of a combustion chamber as iteration variables and takes a bypass airflow state, an injection secondary flow state, the engine thrust and the fuel consumption rate as inspection indexes is used.
For the public understanding, the technical scheme of the invention is explained in detail by a specific embodiment and the accompanying drawings:
taking an AAF fighter in Aircraft engine design written by Mattingly J D as an example, a conventional double-shaft turbofan engine adopted by the fighter is changed into a turbofan engine with an injection nozzle, the structure of the turbofan engine is shown as figure 1, and the injection nozzle consists of a convergent main nozzle and a convergent and divergent outer sleeve. In the figure, a long broken line represents a high-temperature main flow from a core machine, and a short broken line represents an injected low-temperature sub flow; 26. the section 27 is the inlet and outlet of the injection duct, the section 28 is the minimum section of the secondary flow, and the section 29 is the section of the secondary outlet of the spray pipe.
Compared with the conventional engine, the engine with the ejector nozzle has greatly different structures and flow fields, so the improvement needs to be carried out on the basis of a classical parameter cycle analysis method, and the engine with the ejector nozzle can be matched with the flow characteristics of the ejector nozzle.
The input required for the improved parametric loop analysis model is shown in Table 1, wherein the injection bypass ratio α is ej The ratio of the air flow entering the injection duct through the bypass to the flow entering the mixing chamber through the bypass is defined; p is 9 /P 0 And P 29 /P 0 Primary and secondary outlet pressure and ambient back pressure (P), respectively 0 Is the pressure of the engine inlet air stream, the value of which and the ambient back pressure P b Equal), a given 1 indicates a fully expanded state.
TABLE 1 notation of the input variables of the parametric circular analysis model and their meanings
The flow conditions inside the jet nozzle are very complex and have multiple flow states. Referring to the air flow state of the conventional convergent-divergent nozzle, the flow state of the main flow and the secondary flow when the main flow and the secondary flow reach the full expansion is defined as a design state, and the nozzle can obtain the best thrust performance. In this design state, the flow of the air flow inside the nozzle satisfies:
(1) The main flow and the secondary flow are accelerated to an outlet supersonic speed state from an inlet subsonic speed state, and simultaneously, the two gases are in a complete expansion state at the outlet of the spray pipe;
(2) Because the primary flow and the secondary flow at high speed in the injection spray pipe, the momentum and heat exchange are difficult to be fully carried out in a short time, and in order to simplify the calculation, the two gases are supposed not to be mixed in the spray pipe and flow in an isentropic manner;
(3) The static pressure balance is always maintained at the junction of the primary flow and the secondary flow.
The above assumptions can be expressed as:
where M represents the section Mach number, P represents the section static pressure, the subscripts 7, 8, 9 represent the main flow inlet, throat and outlet sections, the subscripts 27, 28, 29 represent the secondary flow inlet, throat and outlet sections, P represents the section static pressure, and b representing ambient backpressure.
The flow of calculation of the ejector nozzle is shown in fig. 2. The known quantities required for the calculation include the primary and secondary flow gas parameters calculated upstream of the nozzle and the relevant input quantities for the parametric loop analysis, the flow conditions being set according to the assumptions above.
The relationship between the air flow passing through the low-pressure turbine and the bypass and the air inflow of the engine is as follows:
the turbofan engine that the fan duct penetrated draws is contained the air current outward and is divided into two strands at the exit, and one and the mixing of core machine air current get into main spray pipe and form the mainstream, and another strand gets into to draw and penetrates the duct and form the secondary flow, and the expression is respectively:
in the formula, alpha ej Is defined as m for injecting bypass ratio 26 And m 16 The ratio of the first to the second.
The relation between the main flow and the secondary flow of the injection nozzle and the air inflow of the engine obtained by the combined formula (2) is as follows:
the static pressure balance of the primary flow and the secondary flow at the junction is assumed in the calculation, namely P 8 ＝P 27 Meanwhile, the sound velocity of the main flow at the outlet of the main nozzle is assumed to be known, so that the ratio P between the total static pressure and the total static pressure of the inlet secondary flow can be calculated t27 /P 27 :
Wherein g (f) 8 ,T t8 ,M 8 ) The oil-gas ratio, the total temperature and the Mach number of the 8-section airflow are used for solving the ratio of the total pressure to the static pressure iteratively.
If the ratio P of the total secondary flow to the static pressure is obtained t27 /P 27 The static pressure at the outlet of the main nozzle is not more than 1, which indicates that the static pressure at the outlet of the main nozzle is too large to inject the bypass secondary flow, and the injection secondary flow is 0 at the moment; if the ratio P of the total secondary flow to the static pressure is obtained t27 /P 27 If the Mach number is larger than 1, the Mach number M can be further iteratively solved 27 ：
M 27 ＝g(f 27 ,T t27 ,P t27 /P t7 ) (6)
Based on the above assumptions about the state of the air flow within the jet nozzle, if M 27 Is greater than1, the injection secondary flow no longer accords with the design state, the static pressure balance condition is not satisfied, and M is set at the moment 27 =1 recalculating the secondary stream; if M is 27 Not more than 1, the primary and secondary flows can satisfy the static pressure balance condition. The flow function of the secondary flow can be obtained according to the gas parameters of the secondary flow, and the flow function of the cross section is defined as follows:
wherein A is a sectional area, γ is a gas specific heat ratio, and R is a gas constant.
As the expansion states of the primary flow and the secondary flow are set in the parameter cycle analysis input quantity, the ratio of the total pressure to the static pressure of the airflow can be calculated, the Mach number of the airflow outlet can be obtained by iterative solution, and the static parameters, the flow speed and the like of the airflow can be further calculated. Taking the mainstream as an example, the expression is as follows:
in the formula, h t And h are the total enthalpy and the static enthalpy of the gas flow, respectively.
The ratio of the total pressure to the flow rate of the main flow and the secondary flow is an important factor for determining the state of the injection secondary flow, and in the input quantity of the parameter cycle analysis model, the fan pressure ratio, the high-pressure compressor pressure ratio, the outlet temperature of the combustion chamber, the total bypass ratio and the injection bypass ratio are key parameters influencing the ratio of the total pressure to the flow rate of the main flow and the secondary flow, wherein the total bypass ratio and the injection bypass ratio are related to the flow rate ratio of the main flow and the secondary flow, the influence rules are similar, and the total bypass ratio is taken as an example for explanation hereinafter.
Fig. 3 is an iterative design flow of the engine core component, in which the 4 model key parameters are used as iterative variables, and initial values of the iterative variables are given according to AAF engine design parameters. The iteration process is roughly divided into 3 parts, firstly, each part from an engine air inlet to a low-pressure turbine is calculated, whether backflow occurs in a culvert is judged according to parameters of a turbine outlet and a culvert inlet, and if abnormal backflow occurs, iteration variables need to be modified; then calculating all parts from the mixing chamber to the injection spray pipe, judging whether the secondary flow reaches the sonic speed according to the airflow parameters at the outlet of the main spray pipe and the outlet of the injection duct, and adjusting an iteration variable to reduce the flow rate of the secondary flow if the secondary flow reaches the sonic speed; and finally, estimating the thrust and oil consumption performance of the engine, wherein the iterative variables are given according to the AAF engine, so that the estimation standard is within 5 percent of the performance gap of the AAF engine.
According to the flow shown in fig. 3, after a plurality of iterative cycles, iterative variable design values satisfying performance requirements can be obtained, and the parameter loop analysis model input amount setting values including the iterative variable design values are shown in table 2. The design point is selected in a high-altitude supersonic flight state, and the afterburner is not opened.
TABLE 2 parameter cycle analysis model input settings
The invention adopts a hot Gas model given by Mattingly J D works of Elements of Propulsion, gas Turbines and rocks Volume, the Gas specific heat capacity can be described as a polynomial of 7 th degree about temperature, and the oil-Gas ratio is adopted for correction so as to improve the accuracy of a calculation result.
In the formula, coefficient A i 、B i (i =1,2 \ 82307; 7) are all known constants, subscript air represents air and corresponds to a value at an oil-gas ratio of 0, subscript prod represents a value at a maximum oil-gas ratio (about 0.0676), and finally, the calculated Cp is a specific heat capacity value at the current oil-gas ratio.
The partial calculation results of the parametric cycle analysis model are shown in Table 3, where F is the combustion chamber fuel-air ratio, F s Is unit thrust, S is unit thrustPower consumption rate.
TABLE 3 results of the calculation of the parameter cycle analysis model
The engine parameter cycle analysis mathematical model is a rubber engine model without given air intake flow, and the cross section of the calculation result is given in a dimensionless parameter mode. The research object of the embodiment is a small bypass ratio turbofan engine, and the air intake flow m of the engine is given for subsequent quasi-one-dimensional design 0 =90 (kg/s), from which a series of critical cross-sectional areas of the exhaust system can be calculated, as shown in table 4.
TABLE 4 critical cross-sectional area values for ejector nozzles
A typical convergent-divergent jet nozzle configuration is shown in fig. 6: d 8 Is the diameter of the outlet of the main nozzle, d s Is the diameter of the throat of the outer sleeve, d 9 The diameter of the outlet of the outer sleeve; theta P For main nozzle convergence half angle, theta C And theta D Respectively the converging and diverging half angles of the outer sleeve.
The geometrical characteristics of the ejector nozzle can be described generally by the following four dimensionless parameters: diameter ratio of throat s /d 8 Throat spacing ratio s/d 8 Outlet diameter ratio d 9 /d 8 Outlet pitch ratio L/d 8 . Wherein d is 8 、d 9 Can be represented by A in Table 3 8 、A 9 And A 29 And directly calculating to obtain that other dimensionless characteristics need to be determined by supplementing the contraction, expansion and half-angle of the spray pipe. The principle of half-angle selection is to make the tail jet flow convert the pressure potential energy into kinetic energy efficiently. If the half angle is too small, the length of the spray pipe is too long, the air flow loss is serious, and the weight of the engine is increased; if the half angle is selected to be too large, the supersonic airflow will be separated from the pipe wall in the expansion section, and eddy current loss is generated.
The specific embodiment designs 3 injection nozzle structural schemes with different structures, and the geometric characteristic parameters of the injection nozzle structural schemes are shown in a table 4. The spray pipe A is a short spray pipe with a larger contraction and expansion half angle, the spray pipe C is a long spray pipe with a smaller contraction and expansion half angle, the contraction and expansion half angle of the spray pipe B is between the spray pipes A and C, and 3 structural schemes are shown in figure 4.
In order to test the accuracy of the parameter cycle analysis model constructed by the method and the performance of the designed spray pipe under various working conditions, a two-dimensional geometric model is established based on commercial CFD calculation software FLUENT according to the injection spray pipe structure scheme. Because the axisymmetric ejector nozzle studied in the invention is rotationally symmetric about the center line, the established two-dimensional geometric model can be used for simulating the flow and heat transfer of a three-dimensional flow field. Fig. 5 is a grid division diagram of the ejector nozzle and the external flow field, the number of grids in the 3 nozzle structure schemes is about 8 ten thousand, and the grid independence verification is carried out. FIG. 6 is a schematic diagram of a simulation calculation domain and boundary conditions, wherein the primary flow inlet and the secondary flow inlet of the injection nozzle are pressure inlets, and the inlet and the outlet of an external flow field are also set as pressure inlets and pressure outlets due to high-altitude supersonic flight at a design point. In the figure P t And T t Respectively represent total temperature and total pressure, P and T respectively represent static temperature and static pressure, and subscripts P, s and f respectively represent a main flow, a secondary flow and an external flow field.
For space reasons, the test results of the parametric loop analysis model are given here as example in the case of scheme B, as shown in table 5. In table V e And M e Respectively, the average speed and Mach number of the outlet of the jet nozzle, T t,e Is the average total temperature of the gas outlet, P e And P b Respectively airflow outlet static pressure and ambient backpressure. Because the parameter cycle analysis is essentially a zero-dimensional mathematical model calculated according to the principle of the aero-thermodynamics, and the CFD flow field calculation is based on a two-dimensional geometric model, the cross-section parameters of the flow field calculation result are subjected to surface integral processing for comparing the calculation results of the two models. Where the flow velocity and mach number are taken as integrals based on mass flow and the other parameters are integrals based on area.
TABLE 5 error of flow field simulation results of parametric cyclic analysis model (PCA) and CFD
As can be seen from the data in the table, the calculation error of the parameter cycle analysis model for the main flow is very low, and the calculation error for the injection flow is slightly higher. The errors mainly come from friction loss of the wall surface, mixing effect between shock wave loss and air flow and the like, and the flow rate of the main flow is obviously larger than that of the injection flow (about 5 times), so that the generated errors have larger influence on the calculation result of the secondary flow. However, in the whole view, the error value calculated by the model is still in an acceptable range, and the model can accurately reflect the flow field characteristics and the performance characteristics of the engine with the ejector nozzle.
The thrust coefficient is an important index for measuring the thrust performance of different spray pipes, is defined as the ratio of the thrust actually generated by the spray pipes to the ideal thrust, and for the injection spray pipes, the ideal thrust comprises two parts of main flow and injection secondary flow, and the calculation expression of the thrust coefficient is as follows:
in the formula, the ideal outlet velocity V of the primary flow and the secondary flow i The method can be calculated by using a one-dimensional entropy flow theory:
in order TO research the thrust performance of the designed injection nozzle under different working conditions, five typical task sections of Takeoff (TO), subsonic cruise (SBC), supersonic cruise (SPC), hovering (HO) and Horizontal Acceleration (HA) are selected TO carry out simulation work. The flight conditions and simulation boundary conditions for each task segment are given in table 7, where the values set for the simulation boundary conditions are derived from the engine complete machine component level model, and are obtained by solving the co-working equation at the corresponding task segment.
TABLE 6 exemplary task Point simulation boundary conditions
FIG. 7 is a graph of thrust and injection coefficients for each mission nozzle. It can be seen from the figure that the thrust coefficients of the 3 types of nozzles are very close, while there is a certain difference in thrust coefficients at different mission segments. The thrust coefficient of the jet nozzle is mainly related to the jet coefficient and the nozzle pressure drop ratio, the jet coefficients of different task sections are all kept near 0.115, and the nozzle pressure drop ratio variation range is very large (3-15).
Fig. 8 is a plot of thrust coefficient of the nozzle as a function of NPR, maintaining a constant bleed coefficient (ω = 0.115). As can be seen, the maximum thrust coefficient (approximately 0.995) is achieved when the nozzle is in the design condition, and the flow is discharged in a nearly fully expanded condition; when the NPR is larger than the design value (P-Q section), the outlet airflow is in an under-expansion state, and the thrust coefficient is slightly reduced; when the NPR is smaller than the design value (P-M section), the thrust coefficient shows a trend of descending first and then ascending as the NPR descends. The main reason for the decrease of the thrust coefficient of the P-O section is the over-expansion of the airflow, and the over-expansion state is more severe the lower the NPR is; the expansion section of the outer sleeve of the O-point spray pipe begins to generate normal shock waves, the normal shock waves in the O-N section pipe continuously move towards the throat, and the thrust coefficient of the spray pipe A in the section is always larger than that of the spray pipes B and C; the normal shock wave of the N-M section disappears, loss caused by the fact that air flow is separated from the wall surface begins to account for the dominant factor, and the C spray pipe has certain advantages due to the fact that the contraction and expansion half angle is small.
By combining the analysis, the thrust coefficient of the jet pipe is insensitive to the contraction and expansion half angle of the injection jet pipe under the high NPR working condition (O-Q section), and the thrust performance of the A jet pipe is excellent under the low NPR working condition (M-O section), so that the thrust performance of the A jet pipe is more excellent in the whole wide NPR change range. Finally, from the simulation result of designing the pumping characteristic and the thrust characteristic of the spray pipe, the spray pipe A can pump enough airflow to meet the requirement of inhibiting infrared radiation, has good thrust performance in a covered wire, has the advantages of small axial length and light weight, and is the spray pipe with the best comprehensive performance in 3 structural schemes.
Claims (3)
1. The design method of the aircraft engine with the ejector nozzle is characterized in that a two-dimensional geometric structure of the ejector nozzle is obtained based on an improved parameter cycle analysis model; in the improved parameter cycle analysis model, the airflow flow inside the injection nozzle meets the following conditions of the Mach number of the main flow and the secondary flow and the static pressure distribution:
where M represents the section Mach number, P represents the section static pressure, the subscripts 7, 8, 9 represent the main flow inlet, throat and outlet sections, the subscripts 27, 28, 29 represent the secondary flow inlet, throat and outlet sections, P represents the section static pressure, and b representing ambient backpressure.
2. The method of claim 1, wherein the model input variables of the improved parametric loop analysis model include a pilot bypass ratio and a ratio of static pressures at the primary and secondary exit ports to ambient backpressure.
3. The method for designing an aircraft engine with an ejector nozzle according to claim 1, wherein in the process of selecting the design value of the engine core component, a multi-stage iterative method is adopted, wherein a bypass ratio, an ejector bypass ratio, a fan pressure ratio, a compressor pressure ratio and a combustion chamber outlet temperature are used as iterative variables, and a bypass airflow state, an ejector secondary flow state, engine thrust and fuel consumption rate are used as inspection indexes.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211432776.1A CN115688287A (en) | 2022-11-16 | 2022-11-16 | Design method of aviation turbofan engine with ejector nozzle |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211432776.1A CN115688287A (en) | 2022-11-16 | 2022-11-16 | Design method of aviation turbofan engine with ejector nozzle |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115688287A true CN115688287A (en) | 2023-02-03 |
Family
ID=85054060
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211432776.1A Pending CN115688287A (en) | 2022-11-16 | 2022-11-16 | Design method of aviation turbofan engine with ejector nozzle |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115688287A (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115879396A (en) * | 2023-03-02 | 2023-03-31 | 中国航发四川燃气涡轮研究院 | Flow one-dimensional pneumatic design method for air inlet front chamber of high-altitude simulation test bed |
| CN116127815A (en) * | 2023-02-28 | 2023-05-16 | 南京航空航天大学 | Modeling method of turbofan engine with injection nozzle |
| CN116595680A (en) * | 2023-05-26 | 2023-08-15 | 中国航发沈阳发动机研究所 | Cross-generation development small-bypass-ratio turbofan engine host and stress application matching method |
| CN116593168A (en) * | 2023-07-14 | 2023-08-15 | 中国人民解放军空军工程大学 | Method and system for evaluating fuel consumption rate of ramjet rotary detonation engine |
-
2022
- 2022-11-16 CN CN202211432776.1A patent/CN115688287A/en active Pending
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116127815A (en) * | 2023-02-28 | 2023-05-16 | 南京航空航天大学 | Modeling method of turbofan engine with injection nozzle |
| CN116127815B (en) * | 2023-02-28 | 2023-11-14 | 南京航空航天大学 | Modeling method of turbofan engine with injection nozzle |
| CN115879396A (en) * | 2023-03-02 | 2023-03-31 | 中国航发四川燃气涡轮研究院 | Flow one-dimensional pneumatic design method for air inlet front chamber of high-altitude simulation test bed |
| CN116595680A (en) * | 2023-05-26 | 2023-08-15 | 中国航发沈阳发动机研究所 | Cross-generation development small-bypass-ratio turbofan engine host and stress application matching method |
| CN116595680B (en) * | 2023-05-26 | 2024-06-11 | 中国航发沈阳发动机研究所 | Cross-generation development small-bypass-ratio turbofan engine host and stress application matching method |
| CN116593168A (en) * | 2023-07-14 | 2023-08-15 | 中国人民解放军空军工程大学 | Method and system for evaluating fuel consumption rate of ramjet rotary detonation engine |
| CN116593168B (en) * | 2023-07-14 | 2023-09-22 | 中国人民解放军空军工程大学 | Method and system for evaluating fuel consumption rate of ramjet rotary detonation engine |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN115688287A (en) | Design method of aviation turbofan engine with ejector nozzle | |
| CN109918839B (en) | Modeling method and infrared radiation prediction method for turbofan engine with pilot nozzle | |
| Presz Jr et al. | Thrust augmentation with mixer/ejector systems | |
| CN111878252B (en) | Air inlet injection nozzle model and turbofan engine model | |
| Zhewen et al. | A multi-fidelity simulation method research on front variable area bypass injector of an adaptive cycle engine | |
| CN112613119A (en) | Multi-geometric-parameter-adjustable air inlet/exhaust/engine integrated aviation propulsion system modeling method | |
| CN109101765A (en) | A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft | |
| Vyvey et al. | Study of an airbreathing variable cycle engine | |
| Tomioka et al. | Sea-level static tests of rocket–ramjet combined cycle engine model | |
| Gamble et al. | Improving off-design nozzle performance using fluidic injection | |
| Yao et al. | Performance analysis of a strut-aided hypersonic scramjet by full-scale IDDES modeling | |
| Li et al. | Numerical investigation of vented plume into a supersonic flow in the early stage of rocket hot separation | |
| CN115618765A (en) | Modeling method of scramjet engine for control based on ASUM format | |
| Sun et al. | Influences of geometric parameters on serpentine nozzles for turbofan | |
| CN114491837A (en) | Design point parameter design method for aircraft engine | |
| CN111523201B (en) | Internal and external flow field coupling iterative calculation method under engine reverse thrust state | |
| Wang et al. | Multi-objective aerodynamic optimization of an axisymmetric variable-geometry inlet with a Mach 5 design point | |
| Nascimento | The selective bleed variable cycle engine | |
| CN116127815B (en) | Modeling method of turbofan engine with injection nozzle | |
| Lin et al. | Effective boundary conditions and numerical method for flow characteristics of aeroengine compressor at high Mach flight | |
| Yarlagadda | Performance analysis of J85 turbojet engine matching thrust with reduced inlet pressure to the compressor | |
| Chen et al. | Modular design framework of an axisymmetric wrap-around thrust-optimized combined nozzle | |
| CN112417658B (en) | Design point parameter cycle modeling method for double-external culvert variable cycle engine | |
| Zhu et al. | Conceptual design and optimization of scramjet engines using the exergy method | |
| Fishbach | Computer simulation of engine systems |
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 |