CN116591857B - Adjustable lobe design method - Google Patents

Adjustable lobe design method Download PDF

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Publication number
CN116591857B
CN116591857B CN202310469072.XA CN202310469072A CN116591857B CN 116591857 B CN116591857 B CN 116591857B CN 202310469072 A CN202310469072 A CN 202310469072A CN 116591857 B CN116591857 B CN 116591857B
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lobe
trough
outlet area
spray pipe
wave
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CN116591857A (en
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施小娟
王丰
俞凡
吉洪湖
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Nanjing University of Aeronautics and Astronautics
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Nanjing University of Aeronautics and Astronautics
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    • 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
    • F02K1/11—Varying effective area of jet pipe or nozzle by means of pivoted eyelids
    • 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

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Nozzles (AREA)

Abstract

The invention relates to the technical field of engine lobe nozzles, in particular to an adjustable lobe design method, which is provided on the basis of the existing adjustable lobe nozzle structure and can meet the requirement of adjusting amplitude of an adjustable lobe in the whole mission flight process of an aircraft to a certain extent. The invention has the advantages of simple design method, simple and clear design flow and easy implementation.

Description

Adjustable lobe design method
Technical Field
The invention belongs to the technical field of engine lobe nozzles, and particularly relates to an adjustable lobe design method.
Background
During the mission of the fighter plane, the relatively high temperature exhaust system cavity and the tail jet are the primary sources of infrared radiation, which results in the fighter plane being the target of detection and attack of the infrared guided weapon. At present, the inhibition of cavity infrared radiation can be realized by a shielding technology, cooling air flow, a stealth coating and other ways, and the inhibition of tail jet infrared radiation is difficult, and researches show that the jet nozzle can enhance the mixing of hot jet and jet air flow, thereby reducing the infrared radiation characteristic of the tail jet.
Compared with the conventional axisymmetric jet nozzle, the lobe jet nozzle has higher jet mixing speed, lower temperature of the tail jet at the outlet after blending and further reduction of infrared radiation of the tail jet. However, the complicated structure of the lobe jet pipe is difficult to realize the adjustment of the outlet area like a conventional axisymmetric jet pipe, which is unfavorable for the good working performance of the engine in a wide speed range. For this purpose, a novel lobe nozzle with adjustable outlet area is designed, and patent number 202310125618X is designed to meet the demands of stealth and flight performance of contemporary fighter aircraft.
Based on the known structure of the adjustable lobe nozzle, how to design the lobes of the main nozzle according to the flight state of the aircraft to determine specific relevant geometric design parameters of the lobes, no relevant published literature exists at present.
Disclosure of Invention
The invention aims to provide a novel parameterized design method for an adjustable lobe, which has the advantages of simple and effective design, parameterized design of the tail edge shape of the adjustable lobe, few design parameters, simple design optimization flow and the like.
The above object of the present invention is achieved by the following technical solutions:
an adjustable lobe design method comprising the steps of:
step one, the area change range of the outlet of the main spray pipe is determined in advance; calculating the total performance of the engine to obtain the outlet area S of the main spray pipe of the engine in different flight mission sections; the main spray pipe is assumed to be the radius R corresponding to the axisymmetric spray pipe p 。
Step two, optimizing and designing the geometrical parameters of the lower lobe of the design point; by utilizing an optimal design method, under the condition that the outlet area of a lobe nozzle is ensured to be equal to the outlet area of a main nozzle required by a typical flight mission section under study, the thrust coefficient, the total pressure recovery coefficient and the infrared suppression effect are comprehensively considered, and an optimal combination A of lobe geometric parameters under a main nozzle design point is studied, wherein the specific lobe geometric parameters comprise the number n of wave crest and wave trough groups and the wave crest height h 1 Height of trough h 2 The peak angle alpha, the trough angle beta, the length L of the lobe along the axis.
Step three, checking the feasibility of the structure; for the main nozzle outlet area of any one flight mission, the lobe geometry parameters are based on the optimal combination A by only changing the peak height h 1 And trough height h 2 Judging whether the peak height h exists 1 And trough height h 2 The formed lobe meets the requirement of the flight mission section on the outlet area of the main spray pipe; if the area requirement of the main spray pipe is met, h corresponding to any flight mission is selected 1 And h 2 In the combination, select (h 1 +h 2 ) min 、(h 1 +h 2 ) max 、(R p +h 1 ) min 、(R p +h 1 ) max 、(R p -h 2 ) min 、(R p -h 2 ) max 。
Continuing to judge whether the following conditions are met:
1.8(h 1 +h 2 ) min ＞(h 1 +h 2 ) max
1.8(R p +h 1 ) min ＞(R p +h 1 ) max ，1.8(R p -h 2 ) min ＞(R p -h 2 ) max
if yes, performing a fourth step of design; if not, reselecting the inferior optimal combination A; the third step is repeated.
Step four, optimizing and designing the geometric parameters of the lower lobe of the non-design point; ensuring that the number n of wave crest and wave trough groups, the wave crest included angle alpha, the wave trough included angle beta and the lobe length L along the axis are equal to corresponding parameter values in the optimal combination A, and researching the peak height h of each task section through an optimization method under the condition that the main spray pipe outlet area of the lobe is ensured to be equal to the main spray pipe outlet area required by other task sections with non-design points under study 1 And trough height h 2 And obtaining an optimal combination set B of h1 and h2 under the comprehensive consideration of the thrust coefficient, the total pressure recovery coefficient and the infrared suppression effect by the two parameters.
Step five, designing the profile of the lobe adjusting piece; the number n of wave crest and wave trough groups, the wave crest included angle alpha, the wave trough included angle beta and the length L of the lobe along the axis select the design parameters of the optimal combination A; (h) 1 +h 2 ) min 、(R p +h 1 ) min And (R) p -h 2 ) min Is the minimum value of the optimal combination set B in all task segments.
The lobe leading edge cross section parameters were chosen as follows:
peak circumferential adjustment blade arc length i=0.95 α×r 0
Trough circumferential adjustment tab arc length j=0.95 β×r 0
R 0 Is the radius of the main flow sleeve;
the lobe trailing edge cross-section parameters were chosen as follows:
peak circumferential conditioner arc length i=0.95 α (R p +h 1 ) min
Trough circumferential adjustment tab arc length j=0.95 beta× (R p -h 2 ) min
Peak radial blade and trough radial blade lengths k=0.95 (h 1 +h 2 ) min 。
As a further improvement of the technology, in the steps (II), (III) and (IV), the outlet area of the lobe nozzle is ensured to meet the requirement of the flight mission section on the outlet area of the main nozzle through the following equation.
(α+β)*n＝2π
π(R p +h 1 ) 2 α/360+π(R p -h 2 ) 2 β/360＝S/n。
The invention has the following advantages:
by adopting the design method of the invention, the lobe nozzle which meets all task sections of the flight can be designed on the basis of the existing lobe structure. The design flow is simple and clear and is easy to master.
Drawings
Fig. 1 is a schematic cross-sectional view of an overall component.
Fig. 2 is a schematic structural view of a main flow sleeve.
FIG. 3 is a schematic distribution of an aircraft aft body, an ejector sleeve, a main flow sleeve, and a conditioner disk.
Fig. 4 is a schematic view of the installation of the actuator cylinder.
Fig. 5 is a schematic view of the structure of the front end of the ejector sleeve.
FIG. 6 is a schematic illustration of the connection of the adjustment tab and the linkage assembly.
Fig. 7 is a schematic view of the installation of the regulator blade.
Fig. 8 is a schematic structural view of the regulating plate.
FIG. 9 is a schematic view of the structure of the trough adjusting sheet.
Fig. 10 is a schematic structural view of a peak adjusting sheet.
Fig. 11 is a schematic illustration of the connection of the actuator cylinder and the actuator ring.
Fig. 12 is a schematic structural view of the connecting rod assembly.
Fig. 13 is a schematic view of lobe nozzle geometry.
FIG. 14 is a schematic diagram of the geometric parameters used with axisymmetric nozzles.
Fig. 15 is an orthogonal representation illustration.
Fig. 16 is a flow chart of a design method.
Fig. 17 is a schematic illustration of lobe configuration labeling.
Reference numerals in the figures: 1. an aircraft aft-body; 2. an injection sleeve; 3. a main flow sleeve; 4. an actuator cylinder; 5. a lobe assembly; 6. a connecting rod assembly; 7. actuating the ring; 8. a first link; 9. a chute; 10. trough adjusting sheets; 11. a wave crest adjusting sheet; 12. a first connection lug; 13. the second connecting support lug; 14. a second link; 15. a third link; 16. a first support; 17. a fourth link; 18. a second support; 19. a fifth link; 20. a sixth link; 21. a third support; 22. a support plate; 23. crest circumferential adjusting pieces; 24. crest radial adjusting pieces; 25. trough circumferential adjustment sheet; 26. trough radial adjusting pieces; 27. and the jet pipe is ejected.
Detailed Description
The following describes in further detail the embodiments of the present invention with reference to the drawings and examples. The following examples or figures are illustrative of the invention and are not intended to limit the scope of the invention.
1. One existing adjustable lobe nozzle structure is disclosed in patent No. 202310125618X.
An existing adjustable lobe spray pipe structure, as shown in fig. 1 and 3, comprises an aircraft rear body 1, an injection sleeve 2, an injection spray pipe 27 and a main flow sleeve 3, and is characterized in that: as shown in fig. 7, the rear end of the main flow sleeve 3 is provided with a plurality of groups of lobe assemblies 5 which are uniformly distributed in the circumferential direction through pin shaft swing; the lobe component 5 comprises two wave trough regulating sheets 10 and two wave crest regulating sheets 11, wherein the two wave crest regulating sheets 11 are mutually overlapped in the circumferential direction to form wave crests, and the two wave trough regulating sheets 10 are mutually overlapped in the circumferential direction to form wave troughs; the peaks and the troughs are alternately distributed in the circumferential direction.
As shown in fig. 7 and 8, adjacent trough adjusting sheets 10 and crest adjusting sheets 11 are connected through the same pin shaft; as shown in fig. 2 and 6, two actuating rings 7 are slidably mounted on the inner wall of the front end of the injection sleeve 2, the actuating ring 7 at the front is respectively connected with the trough adjusting sheets 10 through a plurality of connecting rod assemblies 6, the actuating ring 7 at the rear is respectively connected with the crest adjusting sheets 11 through a plurality of connecting rod assemblies 6, the actuating ring 7 slides along the inner wall of the injection sleeve 2 to drive the corresponding adjusting sheets to rotate around the pin shafts of the adjusting sheets, and the adjacent adjusting sheets always keep close contact in rotation; a plurality of actuating cylinders 4 which are uniformly distributed in the circumferential direction and can control two actuating rings 7 to slide are arranged on the injection sleeve 2.
As shown in fig. 10, the peak adjusting piece 11 is divided into a peak radial adjusting piece 24 and a peak circumferential adjusting piece 23, and as shown in fig. 9, the valley adjusting piece 10 is divided into a valley radial adjusting piece 26 and a valley circumferential adjusting piece 25; as shown in fig. 8, two peak circumferential direction adjustment pieces 23 of the two peak adjustment pieces 11 constituting the peak overlap each other; two trough circumferential adjustment sheets 25 of the two trough adjustment sheets 10 constituting a trough overlap each other; the two wave trough radial direction adjustment pieces 26 of the two wave trough adjustment pieces 10 constituting the wave trough and the adjacent two adjustment pieces of the two wave crest radial direction adjustment pieces 24 of the two wave crest adjustment pieces 11 constituting the wave crest overlap each other.
As shown in fig. 8, the circular arc mating surfaces of two crest circumferential adjustment pieces 23 in two crest adjustment pieces 11 constituting a crest have the same radius and always keep overlapping and seamless when rotating around the respective pin shafts; the circular arc mating surfaces of the two wave trough circumferential adjustment sheets 25 in the two wave trough adjustment sheets 10 forming the wave trough have the same radius, and always keep overlapping and no gap when rotating around the respective pin shafts.
The trough radial adjustment tabs 26 in the trough are parallel to the peak radial adjustment tabs 24 in the adjacent peaks and share the same pin axis.
The mating surfaces of the trough radial adjustment tabs 26 in the trough and the peak radial adjustment tabs 24 in the adjacent peak are flat and seamless.
As shown in fig. 5, the outer circumferential surface of the front end of the injection sleeve 2 is uniformly provided with a plurality of sliding grooves 9 in the circumferential direction, and as shown in fig. 2 and 11, the front end surface of each actuation ring 7 is uniformly and fixedly provided with a plurality of first connection lugs 12 in the circumferential direction; as shown in fig. 2 and 4, the actuating cylinders 4 arranged on the injection sleeve 2 are in one-to-one correspondence with the first connecting lugs 12 arranged on the two actuating rings 7; as shown in fig. 11, a first connecting rod 8 is fixedly installed at the output end of each actuator cylinder 4, and one end of the first connecting rod 8, which is far away from the actuator cylinder 4, passes through a sliding groove 9 formed in the outer wall surface of the front end of the injection sleeve 2 and is hinged with a corresponding first connecting support lug 12.
As shown in fig. 12, the link assembly 6 includes a second connection lug 13, a second link 14, a third link 15, a first support 16, a fourth link 17, a second support 18, a fifth link 19, a sixth link 20, and a third support 21, wherein the second connection lug 13 is fixedly mounted on the actuation ring 7, and one end of the second link 14 is hinged on the second connection lug 13; the first support 16 is fixedly arranged on the main flow sleeve 3, a third connecting rod 15 is arranged on the upper side of the first support 16 in a swinging manner, and one end, far away from the first support 16, of the third connecting rod 15 is hinged with the other end of the second connecting rod 14; the second support 18 is fixedly arranged on the main flow sleeve 3, and a fifth connecting rod 19 is arranged on the upper side of the second support 18 in a swinging manner; one end of the fourth connecting rod 17 is hinged to one end, far away from the second support 18, of the fifth connecting rod 19, and the other end of the fourth connecting rod 17 is hinged to one end, far away from the first support 16, of the third connecting rod 15; the third support 21 is fixedly mounted on the adjusting plate, one end of the sixth connecting rod 20 is hinged to one end of the fifth connecting rod 19, which is far away from the second support 18, and the other end of the sixth connecting rod 20 is hinged to the third support 21.
The third link 15, the fifth link 19, the fourth link 17 and the connection lines between the first support 16 and the second support 18 form a parallelogram structure.
The back end of the main flow sleeve 3 is fixedly provided with a supporting plate 22 which is circumferentially and uniformly distributed, the regulating piece is provided with a supporting lug, and the regulating piece is connected with the supporting plate 22 on the main flow sleeve 3 through a pin shaft by the supporting lug on the regulating piece.
As shown in fig. 13, (a) is a schematic view of the geometry of the main flow sleeve and the main lobe nozzle, and (b) is a schematic view of the cross-sectional parameters of the trailing edge of the lobe; the geometric parameters include the lobe main nozzle length L, the lobe entrance segment radius R0, the radius R1 at the lobe trailing edge trough, the radius R2 at the lobe trailing edge peak, the lobe trailing edge trough circumferential angle α, the lobe trailing edge peak circumferential angle β, the lobe cycle number N (or a fan angle θ=360°/N including one complete peak and one complete trough).
As shown in fig. 14, (a) is a schematic view of geometrical parameters of the main flow sleeve and the axisymmetric main nozzle, and (b) is a schematic view of cross-sectional parameters of the circumferential angle of the nozzle trailing edge being 2θ. The parameters include axisymmetric nozzle length L, nozzle inlet segment radius R0, and nozzle trailing edge radius Rp.
As shown in fig. 17, which is a schematic profile view of the lobe adjusting plate, the geometric parameters include the length L of the lobe nozzle along the axis, the radius R0 of the entrance section of the lobe nozzle, the radius Rp of the exit of the axisymmetric nozzle when the exit area of the lobe nozzle is equal, the peak height h1, the trough height h2, the circumferential angle α of the trough of the lobe trailing edge, the circumferential angle β of the peak of the lobe trailing edge, the arc length I of the leading edge or the trailing edge of the peak circumferential adjusting plate, the arc length J of the leading edge or the trailing edge of the trough circumferential adjusting plate, the length K of the peak radial adjusting plate and the trough radial adjusting plate.
2. The present invention is described in further detail below with respect to the above existing lobe nozzles, in conjunction with the accompanying drawings and the flow chart of the specific implementation. As shown in fig. 16, a schematic diagram of an adjustable lobe design flow is shown, and specific design ideas are as follows:
according to the step (one), the outlet area of the main spray pipe in each flight mission section is determined through the existing engine overall performance calculation program, so that the variation range of the outlet area S of the main spray pipe is determined. Because the different aircraft perform different tasks, the included flight mission segments are different, and basically comprise take-off, subsonic cruise, acceleration, landing and the like.
And (2) optimizing and designing the geometrical parameters of the lower lobe of the design point.
On the premise of ensuring that the outlet area of a main spray pipe of a lobe is equal to the outlet area of the main spray pipe required by a task section of a design point to be researched, the thrust coefficient, the total pressure recovery coefficient and the infrared suppression effect are comprehensively considered, an optimal combination A of the lobe geometric parameters of the task section is researched through an optimization method, and the specific lobe geometric parameters comprise the number n of wave crest and wave trough groups, the wave crest height h1, the wave trough height h2, the wave crest included angle alpha, the wave trough included angle beta and the length L of the lobe along the axis.
The optimization method of the step (two) is exemplified as follows:
1. on the premise of ensuring that the outlet area of the lobe nozzle meets the requirement of a task section of a design point on the outlet area of the main nozzle, adopting an orthogonal experimental method to design a numerical example, and obtaining the optimal combination of the lobe geometric parameters of the task section. Firstly, independent variables in the jet flow are screened out, dimensionless treatment is carried out, and finally, four factors of L/Rp, beta/alpha, n and h2/Rp are selected, wherein Rp is the axisymmetric jet flow outlet radius equal to the outlet area of the lobe jet flow.
2. Based on the characteristic of four factors and multiple levels, an orthogonal table L can be constructed 9 (3 4 ) It indicates that 9 experiments were required and 4 factors were observed, each at a level of 3. The levels of each factor are numbered 1, 2, 3, respectively, and the tests are grouped by a level number of 3, i.e., every third test is grouped, and finally an orthogonal table is generated, as shown in fig. 15. Determining an example sample according to an orthogonal table, carrying out numerical calculation to obtain a sample result, constructing a kriging proxy model, comprehensively considering a thrust coefficient, a total pressure recovery coefficient and an infrared suppression effect, optimizing by utilizing optimizing software to obtain an optimal geometric parameter combination of a task section of a design point, namely, L/Rp, n, beta/alpha and h2/Rp of the optimal combination of the task section of the design point, and further processing to obtain a lobe geometric parameter optimal combination A of the task section.
And (3) according to the second step, the optimal combination of the geometric parameters of the lobe structure of the typical flight mission section is found out. A typical flight mission is generally the most oil-hungry mission, such as subsonic cruise, etc. The optimal combination is found out, so that the thrust performance, the total pressure recovery performance and the infrared suppression effect of the engine can be considered.
And (3) checking the feasibility of the structure.
Structural assessment (1): for the main nozzle outlet area of any one flight mission, the lobe geometry parameters are based on the optimal combination A in the step (two) by only changing the peak height h 1 And trough height h 2 Judging whether the peak height h exists 1 And trough height h 2 The formed lobe meets the requirement of the flight mission section on the outlet area of the main spray pipe.
Structural assessment (2): if the requirement of the main spray pipe outlet area of the structural check (1) is met, h corresponding to any flight mission 1 And h 2 In the combination, select (h 1 +h 2 ) min 、(h 1 +h 2 ) max 、
(R p +h 1 ) min 、(R p +h 1 ) max 、(R p -h 2 ) min 、(R p -h 2 ) max Continuing to judge whether the following conditions are met:
1、1.8(h 1 +h 2 ) min ＞(h 1 +h 2 ) max . Ensure that the adjacent crest radial regulating sheets and trough radial regulating sheets always have the overlapping parts in the lobe regulating process, and the theoretical minimum overlapping degree is equal to 0.2 (h 1 +h 2 ) min No gap exists, and good air tightness is realized.
2、1.8(R p +h 1 ) min ＞(R p +h 1 ) max ，1.8(R p -h 1 ) min ＞(R p -h 2 ) max . Ensuring that the adjacent two wave crest circumferential adjusting sheets always have an overlapping part in the lobe adjusting process, and the theoretical minimum overlapping degree is equal to 0.2 (R p +h 1 ) min No gap exists, and good air tightness is realized. Ensure that two adjacent trough circumferential adjusting sheets always have overlapping parts in the lobe adjusting process, and the theoretical minimum overlapping degree is equal to 0.2 (R p -h 2 ) min No gap exists, and good air tightness is realized.
If yes, carrying out the design of the step (IV); if not, reselecting the inferior optimal combination A; and (3) carrying out structural examination again.
And (fourth) optimizing and designing the geometrical parameters of the lower lobe of the non-design point.
According to the step (II), an optimal combination A is determined, the number n of groups of wave crests and wave troughs, the circumferential included angle alpha of wave crests, the included angle beta of wave troughs and the length L of a lobe along an axis are ensured to be equal to corresponding parameter values in the combination A, and under the premise that the outlet area of a main spray pipe of the lobe is ensured to be equal to the outlet area of a main spray pipe required by other task sections with non-design points under study, the two parameters of the wave crest height h1 and the wave trough height h2 of each task section are studied through an optimization method, so that h1 and h2 combinations which give consideration to the thrust coefficient, the total pressure recovery coefficient and the infrared suppression effect are obtained, and the optimal combination in all the task sections is formed into an optimal combination set B.
And (c) researching an optimal design method of the optimal combination h1 and h2 on the premise of ensuring that the outlet area of the lobe nozzle meets the requirement of the flight mission section on the outlet area of the main nozzle. And h1 and h2 are two parameters which are mutually influenced, h2 is selected as an independent variable, h1 is selected as a dependent variable, a single factor experiment is carried out, the preliminary range of h2 is (0, rp), a plurality of h2 values are selected in the (0, rp) arithmetic, the thrust coefficient, the total pressure recovery coefficient and the infrared suppression effect of the lobe nozzle are comprehensively considered in different combinations, and the optimal h2 and the corresponding h1 are selected.
And (V) designing the profile of the lobe regulating piece.
The design parameters of the optimal combination A determined in the step (II) are selected, wherein the design parameters comprise the number n of wave crest and wave trough groups, the wave crest included angle alpha, the wave trough included angle beta and the length L of a lobe along an axis, the wave crest height h1 and the wave trough height h2 in all the flight mission sections are determined according to the optimal combination set B determined in the step (II) and the step (IV), so that the geometric parameters and the adjusting range of the lobe are already determined, (h1+h2) min, (Rp+h1) min and (Rp-h 2) min are the minimum values of the optimal combination set B in all the mission sections. And the structural examination of the step (III) is carried out. When specifically designing the dimensions of the lobe adjusting plates, it is necessary to reduce the dimensions of the adjusting plates as much as possible, thereby reducing the weight, while ensuring that adjacent adjusting plates always have good tightness during the adjustment process.
The lobe leading edge cross section parameters were chosen as follows:
peak circumferential adjustment blade arc length i=0.95 α×r 0 Trough circumferential adjustment blade arc length is j=0.95 β×r 0 ，R 0 Is the main flow sleeve radius.
The lobe trailing edge cross-section parameters were chosen as follows:
peak circumferential adjustment blade arc length i=0.95 α× (R p +h 1 ) min Trough circumferential adjustment blade arc length j=0.95 β× (R p -h 2 ) min Peak radial blade and trough radial blade lengths k=0.95× (h 1 +h 2 ) min 。
Through the steps, the lobe structure meeting the requirements can be designed.
In the steps (II), (III) and (IV), the outlet area of the adjustable lobe spray pipe in each task section is always ensured to be equal to the axisymmetric main spray pipe outlet area S obtained by performance calculation, so that when the optimal combination of lobe parameters is selected, the flight of a certain task section is not influenced, and the outlet area of the lobe spray pipe is ensured to meet the requirement of the flight task section on the main spray pipe outlet area through the following equation:

Claims (2)

1. an adjustable lobe design method is characterized in that: the method comprises the following steps:
step one, the area change range of the outlet of the main spray pipe is determined in advance; calculating the total performance of the engine to obtain the outlet area S of the main spray pipe of the engine in different flight mission sections; the main spray pipe is assumed to be the radius R corresponding to the axisymmetric spray pipe p ；
Step two, optimizing and designing the geometrical parameters of the lower lobe of the design point; by utilizing an optimal design method, under the condition that the outlet area of a lobe nozzle is ensured to be equal to the outlet area of a main nozzle required by a typical flight mission section under study, the thrust coefficient, the total pressure recovery coefficient and the infrared suppression effect are comprehensively considered, and an optimal combination A of lobe geometric parameters under a main nozzle design point is studied, wherein the specific lobe geometric parameters comprise the number n of wave crest and wave trough groups and the wave crest height h 1 Height of trough h 2 The wave crest included angle alpha, the wave trough included angle beta, and the length L of the lobe along the axis;
step three, checking the feasibility of the structure; for the main nozzle outlet area of any one flight mission, the lobe geometry parameters are based on the optimal combination A by only changing the peak height h 1 And trough height h 2 Judging whether the peak height h exists 1 And trough height h 2 The formed lobe meets the requirement of the flight mission section on the outlet area of the main spray pipe; if the area requirement of the main spray pipe is met, h corresponding to any flight mission is selected 1 And h 2 In the combination, select (h 1 +h 2 ) min 、(h 1 +h 2 ) max 、(R p +h 1 ) min 、(R p +h 1 ) max 、(R p -h 2 ) min 、(R p -h 2 ) max ，
Continuing to judge whether the following conditions are met:
1.8(h 1 +h 2 ) min ＞(h 1 +h 2 ) max
1.8(R p +h 1 ) min ＞(R p +h 1 ) max ，1.8(R p -h 2 ) min ＞(R p -h 2 ) max
if yes, performing a fourth step of design; if not, reselecting the inferior optimal combination A; carrying out the third step again;
step four, optimizing and designing the geometric parameters of the lower lobe of the non-design point; ensuring that the number n of wave crest and wave trough groups, the wave crest included angle alpha, the wave trough included angle beta and the lobe length L along the axis are equal to corresponding parameter values in the optimal combination A, and researching the peak height h of each task section through an optimization method under the condition that the main spray pipe outlet area of the lobe is ensured to be equal to the main spray pipe outlet area required by other task sections with non-design points under study 1 And trough height h 2 Obtaining an optimal combination set B of h1 and h2 under the comprehensive consideration of a thrust coefficient, a total pressure recovery coefficient and an infrared suppression effect by two parameters;
step five, designing the profile of the lobe adjusting piece; the number n of wave crest and wave trough groups, the wave crest included angle alpha, the wave trough included angle beta and the length L of the lobe along the axis select the design parameters of the optimal combination A; (h) 1 +h 2 ) min 、(R p +h 1 ) min And (R) p -h 2 ) min The minimum value of the optimal combination set B in all task segments is set;
the lobe leading edge cross section parameters were chosen as follows:
peak circumferential adjustment blade arc length i=0.95 α×r 0
Trough circumferential adjustment tab arc length j=0.95 β×r 0
R 0 Is the radius of the main flow sleeve;
the lobe trailing edge cross-section parameters were chosen as follows:
peak circumferential conditioner arc length i=0.95 α (R p +h 1 ) min
Trough circumferential adjustment tab arc length j=0.95 beta× (R p -h 2 ) min
Peak radial blade and trough radial blade lengths k=0.95 (h 1 +h 2 ) min 。
2. An adjustable lobe design method as claimed in claim 1 wherein: in the steps (II), (III) and (IV), the outlet area of the lobe nozzle is ensured to meet the requirement of the flight mission section on the outlet area of the main nozzle through the following equation;
(α+β)*n＝2π
π(R p +h 1 ) 2 α/360+π(R p -h 2 ) 2 β/360＝S/n。
CN202310469072.XA 2023-04-27 2023-04-27 Adjustable lobe design method Active CN116591857B (en)

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Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB881974A (en) * 1958-07-21 1961-11-08 Gen Electric Improvements in noise suppression jet propulsion nozzle
RU4109U1 (en) * 1996-12-10 1997-05-16 Акционерное общество открытого типа "ОКБ Сухого" MULTI-PURPOSE HIGH-MANEUVERED SUPERSONIC AIRPLANE, ITS PLANER UNITS, EQUIPMENT AND SYSTEMS
JP2008105039A (en) * 2006-10-24 2008-05-08 Daihen Corp Method for controlling pulse waveform in plasma mig welding
CN105402048A (en) * 2015-11-30 2016-03-16 南京航空航天大学 Low infrared signature lobe injection mixing device used for two-dimensional nozzle outlet
CN115030836A (en) * 2022-06-27 2022-09-09 南京航空航天大学 Wave valve type rear duct ejector with mode adjusting and mixing strengthening functions
CN115199434A (en) * 2022-07-26 2022-10-18 厦门大学 Design method of jet pipe for reducing underwater exhaust noise and retractable lobe jet pipe

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB881974A (en) * 1958-07-21 1961-11-08 Gen Electric Improvements in noise suppression jet propulsion nozzle
RU4109U1 (en) * 1996-12-10 1997-05-16 Акционерное общество открытого типа "ОКБ Сухого" MULTI-PURPOSE HIGH-MANEUVERED SUPERSONIC AIRPLANE, ITS PLANER UNITS, EQUIPMENT AND SYSTEMS
JP2008105039A (en) * 2006-10-24 2008-05-08 Daihen Corp Method for controlling pulse waveform in plasma mig welding
CN105402048A (en) * 2015-11-30 2016-03-16 南京航空航天大学 Low infrared signature lobe injection mixing device used for two-dimensional nozzle outlet
CN115030836A (en) * 2022-06-27 2022-09-09 南京航空航天大学 Wave valve type rear duct ejector with mode adjusting and mixing strengthening functions
CN115199434A (en) * 2022-07-26 2022-10-18 厦门大学 Design method of jet pipe for reducing underwater exhaust noise and retractable lobe jet pipe

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