CN111042949A - Wide-speed-range injection spray pipe integrated with aircraft and design method - Google Patents
Wide-speed-range injection spray pipe integrated with aircraft and design method Download PDFInfo
- Publication number
- CN111042949A CN111042949A CN201911133988.8A CN201911133988A CN111042949A CN 111042949 A CN111042949 A CN 111042949A CN 201911133988 A CN201911133988 A CN 201911133988A CN 111042949 A CN111042949 A CN 111042949A
- Authority
- CN
- China
- Prior art keywords
- nozzle
- flow
- sleeve
- throat
- main
- 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
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/28—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/06—Varying effective area of jet pipe or nozzle
- F02K1/12—Varying effective area of jet pipe or nozzle by means of pivoted flaps
- F02K1/1207—Varying effective area of jet pipe or nozzle by means of pivoted flaps of one series of flaps hinged at their upstream ends on a fixed structure
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/97—Rocket nozzles
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Jet Pumps And Other Pumps (AREA)
Abstract
The invention discloses a wide-speed-range ejector nozzle integrated with an aircraft and a design method thereof, and provides a design process of the ejector nozzle and an optimal value range of each key design parameter, wherein the design process comprises the following steps: the area of the throat of the injection nozzle, the position of the throat of the injection nozzle, the tail edge angle of the expansion section of the injection nozzle, the maximum rotation angle of the auxiliary inlet valve of the third flow path, the length of the main nozzle and the like. The ejector nozzle designed based on the design method can effectively match a turbine-based combined cycle engine in a wide speed domain range, enhances momentum mixing of fluid in a shear layer between a main flow and a secondary flow of the ejector nozzle, reduces bottom resistance of an aircraft in a low Mach number state, meets the adjusting requirement of the outlet area of the ejector nozzle, greatly simplifies an adjusting and actuating mechanism of the ejector nozzle, and greatly improves the performance of the aircraft and the engine in the wide speed domain range.
Description
Technical Field
The invention belongs to the technical field of fluid transmission and control, and relates to a digital switch type inertia hydraulic converter and a working method thereof.
Background
At present, Combined power devices of hypersonic aircrafts mainly comprise two types, namely RBCC (Rocket Based Combined Cycle Engine) and TBCC (Turbine Based Combined Cycle Engine). Compared with RBCC, the TBCC has the characteristics of horizontal take-off and landing, repeated use, good economy, good safety, small technical risk, capability of using conventional carbon-hydrogen fuel and the like, has better engineering application prospect, and is a key technical field for competitive competition of military and strong countries in the world.
As the TBCC exhaust nozzle needs to work in a large falling pressure ratio range from the take-off state to the normal cruising state, the typical exhaust nozzle with the first Mach 0-4 grade has the falling pressure ratio of 60+, the change amplitude of the passing mass flow is large, and the expansion ratio is changed from 2 in the take-off state to 15-20 in the supersonic cruising state, so that the adjustment of the throat and the expansion ratio of the exhaust nozzle needs to be realized through a variable geometry structure. There have been developments including: the adjustable tail-nozzle pipe technical scheme of multiple hinge adjustable structure, urceolus wall translation and various mechanical types such as centre body rotation, though mechanical type regulation can satisfy the regulation of throat area and exit area, but mechanical structure is very complicated, and reliability, life are limited, and need increase a large amount of actuating mechanism, increased aircraft control system's complexity and total weight of structure undoubtedly, for this, urgent need develop the tail-nozzle pipe governing system that the adaptability is stronger, the regulation mode is simpler.
In contrast, the pneumatic jet nozzle adjusting method is simple in structure, easy to implement, free of additional resistance and a feasible technical scheme. The jet nozzle utilizes high-speed jet flow to jet overflow from a boundary layer of an air inlet or gas (secondary flow) from a cooling flow path and an outer bypass from a secondary flow path, and further shear and mix kinetic energy with high-speed gas (primary flow) flowing out of the main jet pipe, so that the kinetic energy of the fluid of the secondary flow is improved, and the fluid of the primary flow and the fluid of the secondary flow jointly flows out of the jet nozzle to improve the thrust. The jet nozzle can also bring other benefits to the design of the aircraft, for example, if the air inlet channel directly overflows to the outside of the aircraft, overflow resistance is generated, but if the scheme of the jet nozzle is adopted, the overflow is jet accelerated after passing through the jet nozzle, certain thrust is generated, and in addition, the regulation requirement of the outlet area can be greatly reduced through the displacement effect of secondary flow on the main flow, so that the outlet actuation system is greatly simplified. However, although the jet nozzle has been used in engineering, the design method of the jet nozzle for the combined engine is not disclosed, and the key technical problem is how to achieve efficient organization of the main and secondary flows or the sufficient mixing between the main and secondary flows and the tertiary flows under a wide range of conditions through pneumatic design and how to avoid the problem of over-expansion of the jet nozzle in a low mach number state.
Disclosure of Invention
In order to solve the problems, the invention discloses a wide-speed-range injection nozzle integrated with an aircraft and a design method thereof, and solves the technical problems of efficiently organizing main and secondary flows or fully mixing the main and secondary flows and tertiary flows under a wide range of conditions through pneumatic design and avoiding over-expansion of a tail nozzle under a low Mach number state.
In order to achieve the aim, the wide-speed-range injection nozzle integrated with the aircraft provided by the invention adopts the following technical scheme:
a wide velocity range ejector nozzle integrated with an aircraft, comprising: the jet nozzle comprises a contraction main nozzle, a secondary flow channel surrounding the contraction main nozzle and a jet nozzle sleeve positioned behind the secondary flow channel; the rear end of the contraction main nozzle is contracted inwards to form a hollow circular truncated cone; the injection nozzle sleeve surrounds the rear part of the shrinkage main nozzle, and the rear section of the injection nozzle sleeve is hinged with a plurality of outlet area adjusting sheets which are annularly arranged; the outlet surrounded by the outlet area adjusting sheet is the final outlet of the injection spray pipe; the outlet area adjusting sheet rotates inwards from a hinge point of the outlet area adjusting sheet and the sleeve of the injection spray pipe to reduce the outlet area; and the rear end of the outer wall of the secondary flow passage is hinged with a third flow passage auxiliary inlet valve, the third flow passage auxiliary inlet valve is positioned between the rear end of the outer wall of the secondary flow passage and the front end of the injection spray pipe sleeve, and when the third flow passage auxiliary inlet valve is closed, the rear end of the third flow passage auxiliary inlet valve is connected with the front end of the injection spray pipe sleeve.
Furthermore, in a low-Mach-number flight state, the third flow path assists the intake valve to be opened, the outlet area adjusting sheet is folded inwards to reduce the outlet area, and at the moment, the jet nozzle generates thrust jointly by the main flow injected in the contracted main nozzle, the secondary flow injected in the secondary flow path and the third flow path fluid from the third flow path auxiliary intake valve; in a supersonic flight state, the third flow path assists the air inlet valve to be closed, the outlet area adjusting sheet is turned outwards to be flat, and the thrust in the injection nozzle is generated by the main flow and the secondary flow.
Further, the third flow path assists the intake valve in turning inward when opened.
Furthermore, a limiter is arranged on the outer side of the outlet area adjusting sheet, so that the outer wall surface of the outlet area adjusting sheet is horizontal to the outer wall surface of the injection nozzle sleeve when the outlet area adjusting sheet is in the maximum expansion state.
The design method of the wide-speed-range injection spray pipe adopts the following technical scheme, and comprises the following steps:
(1) designing parameters of the injection spray pipe are subjected to dimensionless reduction by contracting the diameter D of the inlet of the main spray pipe, so that the designing parameter values are dimensionless quantities, and the origin of coordinates of the injection spray pipe is defined as the position O of the center of a circle of the inlet section of the main spray pipe;
(2) determining the size of the converging primary nozzle of the eductor nozzle, comprising: diameter D of outlet of main nozzlepAnd a main nozzle length L, wherein the outlet diameter DpCalculated based on a flow formula according to the engine flow under the condition that the outlet is caused to reach a critical state; selecting the length L of the main nozzle by setting the dimensionless length of the main nozzle, namely the value range of L/D is 0.20-0.27;
(3) the profile parameters of the induction nozzle sleeve include: guiding deviceJet nozzle sleeve throat diameter DtAnd an axial distance L from the outlet of the main nozzlet-LpThe throat position of sleeve of jet nozzle is determined, and the distance (L) between throat and outlet of main nozzlet-Lp)/DpThe value range of (a) is-0.06-0.04; on the basis of determining the position of the throat, the dimensionless diameter D of the sleeve throat of the injection nozzletThe value range of the/D is 1.09-1.12, and the diameter D of the throat of the sleeve of the injection nozzle is determinedt，；
(4) The profile parameters of the outlet area adjusting sheet comprise that the value range of the trailing edge angle β is 4-8 degrees;
(5) the opening degree of the auxiliary inlet valve of the third flow path is α, the maximum rotation angle α ranges from 25 degrees to 30 degrees if the sub-flow channel is in a backflow state, and the maximum rotation angle α of the auxiliary inlet valve of the third flow path ranges from 5 degrees to 15 degrees if the sub-flow channel is in a forward flow state;
(6) integrating design parameters of the injection nozzle, performing iterative adjustment in the value range, returning to the step (2) to perform value again in the value range of the dimensionless length L/D of the main nozzle if the thrust performance of the injection nozzle does not reach the standard, wherein the value taking method is to increase a value taking interval from the lower limit each time, the value taking interval of the parameter is 0.01, and entering the step (3) if the thrust performance is improved; distance (L) between the throat and the outlet of the main nozzle given in step (3)t-Lp)/DpThe value is re-taken within the value taking range, the starting point is the lower limit, and the value taking interval is 0.01; on the premise of ensuring the improvement of the thrust performance, the dimensionless diameter D of the sleeve throat of the injection nozzle is sequentially settAnd in the design process, if the value of a certain parameter cannot improve the thrust performance, continuing to take values according to the value taking method until the thrust performance is improved, then entering the optimization of the next parameter, and sequentially overlapping the optimized parameters to generate the final pneumatic profile of the ejector nozzle.
Further, the airflow in the secondary flow channel (2) is in a positive direction, namely, downstream, and the airflow is discharged from the right side of the channel; when the flow is found to be in the reverse direction in an experiment or simulation, the maximum rotation angle of the intake valve (3) is assisted by increasing the third flow path until the flow direction of the gas flow in the sub-flow path (2) is in the forward direction.
Further, when the third flow path auxiliary intake valve (3) is closed, the third flow path auxiliary intake valve is completely attached to the injection nozzle sleeve (4), namely, no gap exists between the third flow path auxiliary intake valve (3) and the injection nozzle sleeve (4), and α is always larger than or equal to 0.
Furthermore, the contraction section is designed by adopting a spline curve, the throat part adopts circular arc smooth transition, and the expansion section adopts various straight generatrices or multiple curves; the expansion section bus equation is: a is2*Δx2+a3*Δx3+a4*Δx4Wherein, Δ x and Δ y are respectively the horizontal and vertical coordinate spacing, and Δ y is (2 y-D)t)/(D-Dt),Δx＝(x-Lt)/(Le-Lt)，a2、a3、a4Is the coefficient of an equation, DtIs the diameter of the sleeve throat of the injection nozzle, D is the diameter of the outlet of the injection nozzle, and LtJet nozzle sleeve throat position, LeFor the total length of the injection spray pipe, included angles between tangents of the expansion section bus at the starting point and the end point and the horizontal direction are respectively theta1、θ2Wherein theta1＝0°，θ2Self-setting: given that the desired expansion rate of the gas stream is substantially equivalent before and after, a2＝-0.5tanθ2*(Le-Lt)+2.5，a3＝-1，a4＝0.5tanθ2*(Le-Lt) -0.5; if a large initial expansion rate is desired, a is given2＝0，a3＝4-tanθ2*(Le-Lt)，a4＝tanθ2*(Le-Lt) -3; if a smaller initial expansion rate is desired, a is given2＝6-0.5tanθ2*(Le-Lt)，a3＝-8，a4＝0.5tanθ2*(Le-Lt)+3
Compared with the prior art, the invention has the following beneficial effects:
the ejector nozzle designed based on the design method can effectively match a turbine-based combined cycle engine in a wide speed domain range, enhances momentum mixing of fluid in a shear layer between a main flow and a secondary flow of the ejector nozzle, reduces bottom resistance of an aircraft in a low Mach number state, meets the adjusting requirement of the outlet area of the ejector nozzle, greatly simplifies an adjusting and actuating mechanism of the ejector nozzle, and greatly improves the performance of the aircraft and the engine in the wide speed domain range. The method provides a feasible design method for developing the injection boosting mechanism research and the engine ground test research of the TBCC combined cycle engine.
Drawings
FIG. 1 is a schematic view of the wide velocity range ejector nozzle of the present invention integrated with an aircraft.
Fig. 2 is a pneumatic profile diagram of the wide-speed-range ejector nozzle in a state where Ma is equal to 0.
Fig. 3 is a pneumatic profile diagram of a wide-speed-range ejector nozzle in a state of Ma ═ 4.
Fig. 4 is a flow field structure diagram of the wide-speed-range ejector nozzle obtained through numerical simulation in a state where Ma is 0.
Fig. 5 is a flow field structure diagram of the wide-speed-range ejector nozzle obtained through numerical simulation in a state where Ma is 4.
Detailed Description
Example one
Referring to fig. 1, this embodiment is an embodiment of a wide velocity range ejector nozzle integrated with an aircraft. This wide speed territory draws spouts includes: the jet nozzle comprises a contracted main nozzle 1, a secondary flow channel 2 surrounding the contracted main nozzle 1 and a jet nozzle sleeve 4 positioned behind the secondary flow channel 2. The main spray pipe 1 and the injection spray pipe sleeve 4 are both annular, and the secondary flow channel 2 surrounds the contracted main spray pipe 1 to form an annular channel. The rear end of the contraction main nozzle 1 is contracted inwards to form a hollow circular truncated cone; the injection nozzle sleeve 4 surrounds the rear part of the shrinkage main nozzle, and the rear section of the injection nozzle sleeve is hinged with a plurality of outlet area adjusting blades 7 which are annularly arranged; the outlet surrounded by the outlet area adjusting sheet is the final outlet of the injection spray pipe; the outlet area adjusting sheet rotates inwards from the hinged point of the outlet area adjusting sheet and the sleeve of the injection spray pipe, so that the outlet area is reduced. And the rear end of the outer wall of the secondary flow passage is hinged with a third flow passage auxiliary inlet valve 3, the third flow passage auxiliary inlet valve is positioned between the rear end of the outer wall of the secondary flow passage and the front end of the injection spray pipe sleeve, and when the third flow passage auxiliary inlet valve is closed, the rear end of the third flow passage auxiliary inlet valve is connected with the front end of the injection spray pipe sleeve. And a limiter 5 is arranged on the outer side of the outlet area adjusting sheet 7, so that the outer wall surface 12 of the outlet area adjusting sheet 7 is horizontal to the outer wall surface of the injection nozzle sleeve 4 when the outlet area adjusting sheet is in the maximum expansion state.
In a low-Mach-number flight state, the third flow path auxiliary intake valve 3 is opened, the outlet area adjusting sheet 7 is folded inwards to reduce the outlet area, and at the moment, the jet nozzle generates thrust jointly by the main flow 8 injected in the contracted main nozzle 1, the secondary flow 9 injected in the secondary flow channel 2 and the third flow path 10 fluid from the third flow path auxiliary intake valve 3; in a supersonic flight state, the third flow path assists the closing of the intake valve 3, the outlet area adjusting sheet 7 is turned outwards and leveled, and the thrust in the pilot nozzle is generated by the main flow 8 and the secondary flow 9. The third flow path assists the intake valve 3 to fold inward when opened.
Example two
The second embodiment discloses a pneumatic design method of the wide-speed-range injection nozzle in the first embodiment. Referring to fig. 1, the working condition of this embodiment is the design state (Ma is 4) and the main flow pressure drop ratio is 60, and the detailed implementation steps of this embodiment designed by the method of the present invention will be described below.
(1) Determining the size of the contraction main nozzle 1 by combining the overall demand, wherein the research range of the dimensionless length (L/D) of the contraction main nozzle is 0.20-0.61, and taking 6 sample points in the range to obtain the following research results: with the reduction of the dimensionless length of the contracted main nozzle, the thrust coefficient of the injection nozzle is gradually increased, and the variation range of the corresponding thrust coefficient is 0.949-0.907; the optimal value range of the dimensionless length (L/D) of the contraction main nozzle 1 is 0.20-0.27, and if the injection nozzle is required to have better thrust performance, the lower limit is selected as far as possible; if it is desired that the area adjustment range of the main nozzle is large, the upper limit is taken. The contracted main nozzle 1 of this example takes a dimensionless length of 0.20.
(2) Protection of the sunThe diameter D of the throat of the sleeve 4 of the injection nozzle is determined with the L/D unchangedtAxial distance (L) from the main nozzle outlett-Lp)/DpThe research range is-0.06-0.51, 6 sample points are taken in the range, and the thrust coefficient is obtained through research (L)t-Lp)/DpThe thrust coefficient is increased and gradually reduced, and the variation range of the corresponding thrust coefficient is 0.956-0.950; preferred distance (L) of throat from exit of main nozzlet-Lp)/DpThe value range of (A) is-0.06-0.04, this example (L)t-Lp)/DpTake-0.06.
(3) Maintain L/D, (L)t-Lp)/DpDetermining the dimensionless diameter (D) of the throat of the sleeve 4 of the ejector nozzle without changingtand/D), the research range is 0.98-1.12, 6 sample points are taken in the range, and the thrust coefficient is obtained through research along with DtThe thrust coefficient is gradually increased by increasing the/D, and the variation range of the corresponding thrust coefficient is 0.933-0.935; non-dimensional diameter (D) of throat of sleeve 4 of ejector nozzletThe optimal value range of the/D) is 1.09-1.12, and the throat position of the ejector nozzle sleeve 4 of the embodiment is 1.12.
(4) Maintain L/D, (L)t-Lp)/Dp、DtDetermining β of the tail edge angle of the expansion section of the sleeve 4 of the injection nozzle with the research range of 4-20 degrees, taking 5 sample points in the range, obtaining that the thrust coefficient is gradually reduced along with the increase of β through research, wherein the corresponding thrust coefficient change range is 0.957-0.953, the optimal value range of the tail edge angle β of the expansion section of the injection nozzle 4 is 4-8 degrees, and the tail edge angle of the expansion section of the injection nozzle 4 is 4 degrees.
(5) Keeping the above profile parameters unchanged, determining the maximum rotation angle α of the auxiliary inlet valve 3 of the third flow path, wherein the research range is 5-30 degrees, taking 6 sample points in the range, and obtaining research results, if the secondary flow path 2 is in a backflow state, the thrust coefficient is increased along with the increase of α, the thrust coefficient change range is 0.608-0.732, the preferable value range of α is 25-30 degrees, the lowest limit is taken as far as possible, if the secondary flow path 2 is in a forward flow state, the thrust coefficient is decreased along with the increase of α, the thrust coefficient change range is 0.974-0.759, the preferable value range of α is 5-15 degrees, and the highest limit is taken as far as possible, and the maximum rotation angle α of the auxiliary inlet valve 3 of the third flow path is 5 degrees in the forward flow state, and the maximum rotation angle is 30 degrees in the backflow state.
(6) The expansion section bus of the injection nozzle sleeve 4 is determined, the performance improvement is mainly targeted in the embodiment, the expansion section adopts a plurality of curves, and the expansion section bus equation is as follows: a is2*Δx2+a3*Δx3+a4*Δx4Wherein Δ y ═ 2y-Dt)/(D-Dt),Δx＝(x-Lt)/(Le-Lt) The included angles between the tangent lines of the expansion section generatrix at the starting point and the end point and the horizontal direction are respectively theta1、θ2，θ10 °, θ in the present embodiment2Take 10 degrees and Le-LtWhen the expansion rate of the gas flow is expected to be basically equal to the front and back of the expansion rate of the gas flow, the coefficient of the bus equation can be obtained by a simultaneous bus equation and a tangent equation of a bus end point as follows: a is2＝2.4515，a3＝-1，a4＝-0.4515。
(7) Integrating the optimal values of the design parameters in the process to generate a profile of the ejector nozzle, wherein the thrust coefficient of the ejector nozzle in the embodiment under the design state (Ma is 4) is 0.961.
The pneumatic scheme of the ejector nozzle based on the invention is shown in fig. 2 and 3, and the flow structure of the ejector nozzle is obtained by using the numerical simulation result, as shown in fig. 4 and 5. In the numerical simulation, when Ma is 0, the main flow drop pressure ratio is 2.16, and the sub flow drop pressure ratio is 0.987; when Ma is 4, the major flow drop pressure ratio is 60.7, and the minor flow drop pressure ratio is 7.1. It can be seen from fig. 4 and 5 that there is no large-scale backflow region in the flow field, the flow field of the primary and secondary flows is uniform, and the free shear layer is fully developed. The thrust coefficients corresponding to the two states are 0.947 and 0.961 respectively. The above results show that the design method of the present invention achieves the intended goal and is feasible.
In addition, the present invention has many specific implementations and ways, and the above description is only a preferred embodiment of the present invention. It should be noted that, for those skilled in the art, without departing from the principle of the present invention, several improvements and modifications can be made, and these improvements and modifications should also be construed as the protection scope of the present invention.
Claims (8)
1. The utility model provides a wide speed region draws and penetrates spray tube with aircraft integration which characterized in that includes: the jet nozzle comprises a contraction main nozzle (1), a secondary flow channel (2) surrounding the contraction main nozzle (1), and a jet nozzle sleeve (4) positioned behind the secondary flow channel (2); the rear end of the contraction main nozzle (1) is contracted inwards to form a hollow circular truncated cone; the injection nozzle sleeve (4) surrounds the rear part of the shrinkage main nozzle, and the rear section of the injection nozzle sleeve is hinged with a plurality of outlet area adjusting sheets (7) which are annularly arranged; the outlet surrounded by the outlet area adjusting sheet is the final outlet of the injection spray pipe; the outlet area adjusting sheet rotates inwards from a hinge point of the outlet area adjusting sheet and the sleeve of the injection spray pipe to reduce the outlet area;
and the rear end of the outer wall of the secondary flow passage is hinged with a third flow passage auxiliary inlet valve (3), the third flow passage auxiliary inlet valve is positioned between the rear end of the outer wall of the secondary flow passage and the front end of the injection spray pipe sleeve, and when the third flow passage auxiliary inlet valve is closed, the rear end of the third flow passage auxiliary inlet valve is connected with the front end of the injection spray pipe sleeve.
2. The wide-velocity-range ejector nozzle according to claim 1, wherein in a low-mach-number flight state, the third flow path auxiliary intake valve (3) is opened, the outlet area adjusting piece (7) is folded inwards to reduce the outlet area, and the ejector nozzle generates thrust jointly by the main flow (8) injected in the contracted main nozzle (1), the secondary flow (9) injected in the secondary flow path (2) and the third flow path (10) from the third flow path auxiliary intake valve (3); in a supersonic flight state, the third flow path assists the closing of the air inlet valve (3), the outlet area adjusting sheet (7) is turned outwards and leveled, and the thrust in the injection nozzle is generated by the main flow (8) and the secondary flow (9).
3. The wide velocity range jet nozzle according to claim 1 or 2, wherein the third flow path assists the inward fold when the inlet valve (3) is open.
4. The wide velocity range ejector nozzle according to claim 1 or 2, wherein the stopper (5) is mounted outside the outlet area adjusting piece (7) so that the outer wall surface (12) of the outlet area adjusting piece (7) is horizontal to the outer wall surface of the ejector nozzle sleeve (4) when the outlet area adjusting piece is at maximum expansion.
5. A method of designing a wide velocity range ejector nozzle according to any one of claims 1 to 3, comprising the steps of:
(1) designing parameters of the injection spray pipe are subjected to dimensionless reduction by contracting the diameter D of the inlet of the main spray pipe, so that the designing parameter values are dimensionless quantities, and the origin of coordinates of the injection spray pipe is defined as the position O of the center of a circle of the inlet section of the main spray pipe;
(2) determining the size of a convergent main nozzle (1) of a pilot nozzle, comprising: diameter D of outlet of main nozzlepAnd a main nozzle length L, wherein the outlet diameter DpCalculated based on a flow formula according to the engine flow under the condition that the outlet is caused to reach a critical state; selecting the length L of the main nozzle by setting the dimensionless length of the main nozzle, namely the value range of L/D is 0.20-0.27;
(3) the profile parameters of the ejector nozzle sleeve (4) comprise: diameter D of throat of sleeve (4) of injection nozzletAnd an axial distance L from the outlet of the main nozzlet-LpThe throat position of the sleeve (4) of the ejector nozzle is determined, and the distance (L) between the throat and the outlet of the main nozzlet-Lp)/DpThe value range of (a) is-0.06-0.04; on the basis of determining the position of the throat, the dimensionless diameter D of the throat is determined by the sleeve (4) of the injection nozzletThe value range of/D is 1.09-1.12, and the diameter D of the throat of the ejector nozzle sleeve (4) is determinedt，；
(4) The profile parameters of the outlet area adjusting sheet (7) comprise that the value range of the trailing edge angle β is 4-8 degrees;
(5) the opening degree of the third flow path auxiliary inlet valve (3) is α degrees at the maximum rotation angle, if the sub-flow channel is in a backflow state, the value range of α degrees at the maximum rotation angle is 25-30 degrees, and if the sub-flow channel (2) is in a forward flow state, the value range of α degrees at the maximum rotation angle of the third flow path auxiliary inlet valve (3) is 5-15 degrees;
(6) integrated ejector nozzleThe design parameters are iteratively adjusted in the value range, if the thrust performance of the injection nozzle does not reach the standard, the step (2) is returned to and the value is re-taken in the value range of the dimensionless length (L/D) of the main nozzle, the value taking method is that a value taking interval is increased from the lower limit every time, the value taking interval of the parameters is 0.01, and if the thrust performance is improved, the step (3) is carried out; distance (L) between the throat and the outlet of the main nozzle given in step (3)t-Lp)/DpThe value is re-taken within the value taking range, the starting point is the lower limit, and the value taking interval is 0.01; on the premise of ensuring the thrust performance to be improved, the dimensionless diameter D of the throat of the sleeve (4) of the injection nozzle is sequentially settAnd in the design process, if the value of a certain parameter cannot improve the thrust performance, continuing to take values according to the value taking method until the thrust performance is improved, then entering the optimization of the next parameter, and sequentially overlapping the optimized parameters to generate the final pneumatic profile of the ejector nozzle.
6. The design method according to claim 5, characterized in that the direction of the gas flow in the secondary flow channel (2) is positive, i.e. downstream, and the gas flow is discharged from the right side of the channel; when the flow is found to be in the reverse direction in an experiment or simulation, the maximum rotation angle of the intake valve (3) is assisted by increasing the third flow path until the flow direction of the gas flow in the sub-flow path (2) is in the forward direction.
7. The design method according to claim 5, characterized in that the third flow path auxiliary intake valve (3) completely fits the jet nozzle sleeve (4) when closed, i.e. no gap exists between the third flow path auxiliary intake valve (3) and the jet nozzle sleeve (4), and α is always greater than or equal to 0.
8. The design method of claim 5, wherein the contraction section is designed by spline curve, the throat is smoothly transited by circular arc, and the expansion section is designed by multiple sectionsPlanting a straight bus or a multiple curve; the expansion section bus equation is: a is2*Δx2+a3*Δx3+a4*Δx4Wherein, Δ x and Δ y are respectively the horizontal and vertical coordinate spacing, and Δ y is (2 y-D)t)/(D-Dt),Δx＝(x-Lt)/(Le-Lt)，a2、a3、a4Is the coefficient of an equation, DtIs the diameter of the sleeve throat of the injection nozzle, D is the diameter of the outlet of the injection nozzle, and LtJet nozzle sleeve throat position, LeFor the total length of the injection spray pipe, included angles between tangents of the expansion section bus at the starting point and the end point and the horizontal direction are respectively theta1、θ2Wherein theta1＝0°，θ2Self-setting: given that the desired expansion rate of the gas stream is substantially equivalent before and after, a2＝-0.5tanθ2*(Le-Lt)+2.5，a3＝-1，a4＝0.5tanθ2*(Le-Lt) -0.5; if a large initial expansion rate is desired, a is given2＝0，a3＝4-tanθ2*(Le-Lt)，a4＝tanθ2*(Le-Lt) -3; if a smaller initial expansion rate is desired, a is given2＝6-0.5tanθ2*(Le-Lt)，a3＝-8，a4＝0.5tanθ2*(Le-Lt)+3。
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911133988.8A CN111042949B (en) | 2019-11-19 | 2019-11-19 | Wide-speed-range injection spray pipe integrated with aircraft and design method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911133988.8A CN111042949B (en) | 2019-11-19 | 2019-11-19 | Wide-speed-range injection spray pipe integrated with aircraft and design method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN111042949A true CN111042949A (en) | 2020-04-21 |
| CN111042949B CN111042949B (en) | 2021-06-29 |
Family
ID=70231897
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201911133988.8A Active CN111042949B (en) | 2019-11-19 | 2019-11-19 | Wide-speed-range injection spray pipe integrated with aircraft and design method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN111042949B (en) |
Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112035952A (en) * | 2020-08-21 | 2020-12-04 | 南京航空航天大学 | Design method of ejector nozzle experimental device for simulating outflow of aircraft |
| CN112179605A (en) * | 2020-08-21 | 2021-01-05 | 南京航空航天大学 | Ejector nozzle experimental device for simulating outflow of aircraft |
| CN113032893A (en) * | 2021-02-22 | 2021-06-25 | 南京航空航天大学 | Design method of ejector nozzle device for simulating subsonic/transonic outflow |
| CN113090411A (en) * | 2021-04-23 | 2021-07-09 | 西北工业大学 | Three-duct S-shaped bent spray pipe with turbulence rib-air film cooling structure |
| CN113236442A (en) * | 2021-06-04 | 2021-08-10 | 中国人民解放军国防科技大学 | Wide-speed-range rocket ramjet and rocket |
| CN114021298A (en) * | 2022-01-06 | 2022-02-08 | 中国空气动力研究与发展中心计算空气动力研究所 | Shrinkage expansion spray pipe marking die design method suitable for jet flow noise research |
| CN114109644A (en) * | 2021-10-09 | 2022-03-01 | 中国航发贵阳发动机设计研究所 | Method for adjusting area of exit throat of special-shaped fixed spray pipe of aircraft engine |
| CN114329847A (en) * | 2022-01-06 | 2022-04-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Design method for 'double-bell' shrinkage expansion spray pipe marking die |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6164059A (en) * | 1995-11-17 | 2000-12-26 | United Technologies Corporation | Multi-expansion ejector nozzle with diverging walls |
| CN102635578A (en) * | 2011-12-23 | 2012-08-15 | 南京航空航天大学 | Multilevel lobed nozzle ejector with secondary-fluid sucking function |
| CN105730701A (en) * | 2016-02-18 | 2016-07-06 | 江西洪都航空工业集团有限责任公司 | Secondary flow system capable of changing secondary flow inlet area |
| CN108087147A (en) * | 2017-11-29 | 2018-05-29 | 中国直升机设计研究所 | A kind of non-homogeneous injection ejector exhaust pipe |
| CN108999725A (en) * | 2018-07-19 | 2018-12-14 | 北京航空航天大学 | A kind of nozzles with injector of the double bell injection casings of band |
-
2019
- 2019-11-19 CN CN201911133988.8A patent/CN111042949B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6164059A (en) * | 1995-11-17 | 2000-12-26 | United Technologies Corporation | Multi-expansion ejector nozzle with diverging walls |
| CN102635578A (en) * | 2011-12-23 | 2012-08-15 | 南京航空航天大学 | Multilevel lobed nozzle ejector with secondary-fluid sucking function |
| CN105730701A (en) * | 2016-02-18 | 2016-07-06 | 江西洪都航空工业集团有限责任公司 | Secondary flow system capable of changing secondary flow inlet area |
| CN108087147A (en) * | 2017-11-29 | 2018-05-29 | 中国直升机设计研究所 | A kind of non-homogeneous injection ejector exhaust pipe |
| CN108999725A (en) * | 2018-07-19 | 2018-12-14 | 北京航空航天大学 | A kind of nozzles with injector of the double bell injection casings of band |
Non-Patent Citations (1)
| Title |
|---|
| LEONARD E.STITT: "Exhaust Nozzles for Propulsion Systems with Emphasis on Supersonic Cruise Aircraft", 《NASA REFERENCE PUBLICATION 1234》 * |
Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112179605B (en) * | 2020-08-21 | 2021-10-01 | 南京航空航天大学 | Ejector nozzle experimental device for simulating outflow of aircraft |
| CN112179605A (en) * | 2020-08-21 | 2021-01-05 | 南京航空航天大学 | Ejector nozzle experimental device for simulating outflow of aircraft |
| CN112035952A (en) * | 2020-08-21 | 2020-12-04 | 南京航空航天大学 | Design method of ejector nozzle experimental device for simulating outflow of aircraft |
| CN112035952B (en) * | 2020-08-21 | 2021-07-27 | 南京航空航天大学 | Design method of ejector nozzle experimental device for simulating outflow of aircraft |
| CN113032893A (en) * | 2021-02-22 | 2021-06-25 | 南京航空航天大学 | Design method of ejector nozzle device for simulating subsonic/transonic outflow |
| CN113032893B (en) * | 2021-02-22 | 2024-03-29 | 南京航空航天大学 | Design method of jet nozzle device for simulating sub/transonic outflow |
| CN113090411A (en) * | 2021-04-23 | 2021-07-09 | 西北工业大学 | Three-duct S-shaped bent spray pipe with turbulence rib-air film cooling structure |
| CN113090411B (en) * | 2021-04-23 | 2023-02-10 | 西北工业大学 | Three-duct S-shaped bent spray pipe with turbulence rib-air film cooling structure |
| CN113236442B (en) * | 2021-06-04 | 2022-02-22 | 中国人民解放军国防科技大学 | Wide-speed-range rocket ramjet and rocket |
| CN113236442A (en) * | 2021-06-04 | 2021-08-10 | 中国人民解放军国防科技大学 | Wide-speed-range rocket ramjet and rocket |
| CN114109644A (en) * | 2021-10-09 | 2022-03-01 | 中国航发贵阳发动机设计研究所 | Method for adjusting area of exit throat of special-shaped fixed spray pipe of aircraft engine |
| CN114021298A (en) * | 2022-01-06 | 2022-02-08 | 中国空气动力研究与发展中心计算空气动力研究所 | Shrinkage expansion spray pipe marking die design method suitable for jet flow noise research |
| CN114021298B (en) * | 2022-01-06 | 2022-03-15 | 中国空气动力研究与发展中心计算空气动力研究所 | Shrinkage expansion spray pipe marking die design method suitable for jet flow noise research |
| CN114329847A (en) * | 2022-01-06 | 2022-04-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Design method for 'double-bell' shrinkage expansion spray pipe marking die |
| CN114329847B (en) * | 2022-01-06 | 2023-03-21 | 中国空气动力研究与发展中心计算空气动力研究所 | Design method for 'double-bell' shrinkage expansion spray pipe marking die |
Also Published As
| Publication number | Publication date |
|---|---|
| CN111042949B (en) | 2021-06-29 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111042949B (en) | Wide-speed-range injection spray pipe integrated with aircraft and design method | |
| EP1718857B1 (en) | Integrated air inlet system for multi-propulsion aircraft engines | |
| US8393158B2 (en) | Low shock strength inlet | |
| US8166768B2 (en) | Systems and methods for passively directing aircraft engine nozzle flows | |
| US8596037B2 (en) | Nacelle with a displacement device for aircraft jet engine and aircraft including such nacelle | |
| US7818958B2 (en) | Jet engine nacelle for an aircraft and aircraft comprising such a nacelle | |
| EP1995443B1 (en) | Method and valve assembly for regulating fluid flow through a turbine engine | |
| CN110284994B (en) | Parallel thrust vector exhaust system based on throat offset type pneumatic vector spray pipe | |
| US20100037587A1 (en) | Turbojet Nacelle Equipped With Means For Reducing The Noise Produced By Said Turbojet | |
| CN108561244B (en) | The three power combination engine design methods that a kind of super burn and sub- combustion combustion chamber coexist | |
| JPH0350100A (en) | Hybrid laminar flow nacelle | |
| US20180094605A1 (en) | Turbofan engine for a civil supersonic aircraft | |
| CN112035952B (en) | Design method of ejector nozzle experimental device for simulating outflow of aircraft | |
| US20090107108A1 (en) | Nacell for bypass engine with high bypass ratio | |
| CN106014684A (en) | Combined flow control method and structure for improving SERN for TBCC | |
| US10443539B2 (en) | Hybrid exhaust nozzle | |
| CN101392686A (en) | Internal waverider hypersonic inlet and design method for giving attention to internal and external flow performance | |
| EP2865874B1 (en) | Turbofan engine with passive thrust vectoring | |
| CN103987948B (en) | Effuser device and manufacture method thereof | |
| CN105464838B (en) | Method and apparatus for being deflected by dynamicthrust guiding and plume | |
| CN201301753Y (en) | Inner wave rider type air inlet channel taking internal and external flow performance into consideration | |
| EP1710426A2 (en) | Combi-supersonic-adjusting-nozzle | |
| Conners et al. | The feasibility of high-flow nacelle bypass for low sonic boom propulsion system design | |
| CN103890349B (en) | Atmospheric gas inlet unit and its manufacture method | |
| CN112179605B (en) | Ejector nozzle experimental device for simulating outflow of aircraft |
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 |