CN117465664A - Aircraft integrating fluid thrust vectoring nozzle and wing - Google Patents
Aircraft integrating fluid thrust vectoring nozzle and wing Download PDFInfo
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- CN117465664A CN117465664A CN202311292108.8A CN202311292108A CN117465664A CN 117465664 A CN117465664 A CN 117465664A CN 202311292108 A CN202311292108 A CN 202311292108A CN 117465664 A CN117465664 A CN 117465664A
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- wing
- integrated
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- 239000012530 fluid Substances 0.000 title claims abstract description 39
- 230000002706 hydrostatic effect Effects 0.000 claims description 15
- 238000005452 bending Methods 0.000 claims description 8
- 238000004891 communication Methods 0.000 claims description 5
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- 238000000034 method Methods 0.000 abstract description 9
- 230000008569 process Effects 0.000 abstract description 3
- 230000004044 response Effects 0.000 abstract description 3
- 239000013598 vector Substances 0.000 description 11
- 238000013461 design Methods 0.000 description 8
- 230000000694 effects Effects 0.000 description 7
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- 230000005855 radiation Effects 0.000 description 7
- 238000005516 engineering process Methods 0.000 description 6
- 238000010586 diagram Methods 0.000 description 4
- 239000007921 spray Substances 0.000 description 4
- 238000012360 testing method Methods 0.000 description 4
- 230000009467 reduction Effects 0.000 description 3
- 230000008859 change Effects 0.000 description 2
- 230000008878 coupling Effects 0.000 description 2
- 238000010168 coupling process Methods 0.000 description 2
- 238000005859 coupling reaction Methods 0.000 description 2
- 238000001514 detection method Methods 0.000 description 2
- 230000005484 gravity Effects 0.000 description 2
- 238000002347 injection Methods 0.000 description 2
- 239000007924 injection Substances 0.000 description 2
- 238000004088 simulation Methods 0.000 description 2
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 description 1
- 230000009471 action Effects 0.000 description 1
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- 238000004401 flow injection analysis Methods 0.000 description 1
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- 230000002401 inhibitory effect Effects 0.000 description 1
- 238000013507 mapping Methods 0.000 description 1
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C21/00—Influencing air flow over aircraft surfaces by affecting boundary layer flow
- B64C21/02—Influencing air flow over aircraft surfaces by affecting boundary layer flow by use of slot, ducts, porous areas or the like
- B64C21/08—Influencing air flow over aircraft surfaces by affecting boundary layer flow by use of slot, ducts, porous areas or the like adjustable
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C15/00—Attitude, flight direction, or altitude control by jet reaction
- B64C15/02—Attitude, flight direction, or altitude control by jet reaction the jets being propulsion jets
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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
- F02K1/002—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto with means to modify the direction of thrust vector
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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
- F02K1/78—Other construction of jet pipes
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention discloses an aircraft with integrated fluid thrust vectoring nozzle and wing, comprising an aircraft integrated wing section, an engine tail nozzle and a fluid thrust vectoring nozzle; the engine is embedded in the flying integrated wing section, and the tail nozzle of the engine is preset to be upwards deviated. According to the aircraft with the integrated fluid thrust vectoring nozzle and the wing, the thickness of the trailing edge of the wing is effectively reduced, so that the rear body resistance in the flight process is reduced; the structure is light in weight, few in movable parts and quick in deflection response, and can provide controllable pitching moment in the cruising attack angle of the aircraft; the lift coefficient and the control moment are improved; the infrared stealth performance of the aircraft is enhanced.
Description
Technical Field
The invention belongs to the technical field of vector control of aircrafts, and particularly relates to an aircraft with a fluid thrust vector nozzle and an airfoil.
Background
Thrust vector technology has been widely used to improve the maneuverability and agility of new generation fighters. The thrust vectors are classified into two types, mechanical and fluid, according to structural modes and deflection mechanisms. The mechanical thrust vector changes the jet flow direction through the mechanical profile deflection of the tail part of the jet pipe, the design is mature, and the jet flow direction is widely applied to modern fighters; however, the structure is heavy, the number of movable parts is large, and the deflection response is slow. The fluid thrust vectoring nozzle does not need to change the shape of the nozzle, and changes the direction of the jet by means of flow control methods such as jet injection and the like, so that the fluid thrust vectoring nozzle has a relatively simple structure and fewer movable parts.
The existing fluid thrust vectoring nozzle is mainly installed at the tail of an aircraft to provide control moment, and has less design consideration on the integrated design of the aircraft. Researchers such as Han Jiexing install rectangular passive fluid thrust vector nozzles on the trailing edge of the wing, and verify the effect of jet flow on the lift and drag increase of the wing shape through wind tunnel tests. However, the thicker rectangular passive fluid thrust vectoring nozzle thickness h1 (as shown in fig. 1) can cause greater aft body drag on the aircraft, and a distance from the actual application.
Most of the surface experiments and numerical simulations for the study of fluid vectoring nozzles showed that: the free incoming flow can cause the pressure near the outlet of the spray pipe to change and act on the boundary layer of the inner wall surface of the spray pipe, thereby further influencing the flow field in the spray pipe. The annular quantity control technology is used as one of the flow control modes, and the annular quantity is changed by forming a coanda effect by blowing air at the trailing edge of the wing, so that the lift-increasing drag reduction and attitude control of the aircraft are realized, and the method has the advantages of simplicity in use, light weight, easiness in realization and the like. The annular quantity control technology has wide application prospects in the aspects of performance enhancement of the aircraft, such as stall attack angle improvement, flow separation delay, lift increase drag reduction and the like, and is also increasingly valued in innovative flight control.
Disclosure of Invention
The invention aims to solve the technical problem of providing an aircraft with integrated fluid thrust vectoring nozzle and wing, which can reduce the tail thickness of wing airfoil and reduce the pressure difference resistance of the aircraft while providing the attitude control moment of the aircraft.
In order to achieve the above purpose, the present invention adopts the following technical scheme:
an aircraft with integrated fluid thrust vectoring nozzle and wing comprises an integrated aerofoil, an engine tail nozzle and a fluid thrust vectoring nozzle. The engine is embedded in the flying integrated wing section, and the tail nozzle of the engine is preset to be upwards deviated.
Further, the engine tail nozzle comprises a round-square rectifying section and a square bending deflection section which are sequentially connected.
Still further, the aspect ratio of the rectifying section depends on the number of engines, the jet outlet area and the flow size of the working point jet outlet.
Further, the fluid thrust vectoring nozzle comprises a plurality of sections of nozzles, and each section of nozzle comprises a deflection wall plate, a deflection control hole, a deflection control valve and a deflection hydrostatic cavity.
Further, the wall surface of the deflection wall plate is a plane rectangle, and the aspect ratio of the wall surface is determined by the jet expansion angle and the space position of the shear layer, and is obtained through experiments.
Still further, the sum of the areas of all deflection-control holes on the same wall surface is not more than 10% of the area of the wall surface.
Further, the deflection control hole is arranged at the front end of the deflection wall plate; the diameter and number of deflection control holes are dependent on the size of the secondary flow rate required.
Further, the deflection hydrostatic cavity is used for connecting the deflection control hole with the outside atmosphere.
Further, the deflection control valve is used for changing the communication degree of the deflection hydrostatic cavity.
Further, the deflection control valve comprises a linear limit groove, a valve plate and a driving device; the linear limit groove is used for limiting the moving track of the valve plate, so that the valve plate can only move along a straight line; the driving device is used for controlling the valve plate to be at the position on the linear limit groove. The opening degree of the deflection control valve between the deflection hydrostatic cavity and the atmosphere can be changed by changing the running position of the valve plate on the linear limit groove, so that the communication degree between the deflection hydrostatic cavity and the outside atmosphere and the transported flow are adjusted, and finally the coanda degree between jet flow and the wall surface is changed, so that the jet flow direction is changed in a linear and controllable manner.
The invention discloses an aircraft integrating a fluid thrust vectoring nozzle and a wing, which has the following beneficial effects:
compared with the design that the traditional nozzle is positioned at the tail end of the wing, the technical scheme of the invention reduces the thickness of the rear body of the wing trailing edge by leading the nozzle, and effectively reduces the thickness of the wing trailing edge, thereby reducing the rear body resistance in the flying process;
the technical scheme of the invention adopts the fluid type thrust vector as a thrust vector deflection control device, has light structure weight, few movable parts and quick deflection response, and can provide controllable pitching moment in the cruising attack angle of the aircraft.
According to the technical scheme, through mixing between jet flow and outflow flow of the engine, the temperature of the jet flow of the engine can be reduced, and meanwhile, the jet flow is shielded by the multi-section deflection wall plates from the lower part, so that the infrared stealth performance of the aircraft is enhanced.
The technical scheme of the invention adopts an internal and external flow coupling integrated design, provides a control moment for a normal component of thrust vector force, and can be simultaneously used as a propulsion device and a flow control device to generate a superannular effect and improve a lift coefficient and a control moment at the same time; the method has application prospect in the fields of short-distance take-off and landing flight, high maneuvering fighter and the like.
Drawings
FIG. 1 is a conventional fluid thrust vector tube and wing integrated design;
FIG. 2 is a schematic illustration of the aft body thickness of an aircraft incorporating a fluid thrust vectoring nozzle and airfoil of the present invention;
FIG. 3 is a schematic view of the overall structure of an aircraft with integrated fluid thrust vectoring nozzle and wing according to the present invention;
FIG. 4 is a schematic view of a portion of the structure of a fluid thrust vectoring nozzle of the present invention;
FIG. 5 is a schematic illustration of the opening of a two-stage deflection control valve in one embodiment of the present invention;
FIG. 6 is a state diagram of the aircraft in the open and closed states shown by the yaw control valve of FIG. 5;
FIG. 7 is a schematic illustration of one section of a deflection control valve being open and another section of the deflection control valve being closed in one embodiment of the present invention;
FIG. 8 is a state diagram of the aircraft in the yaw control valve of FIG. 7 in an open and closed state;
FIG. 9 is a schematic illustration of two-stage deflection control valves each closed in one embodiment of the present invention;
FIG. 10 is a state diagram of the aircraft in the yaw control valve of FIG. 9 in an open and closed state;
FIG. 11 is a diagram of pitch test results for an aircraft wing having an angle of attack in the range of-4 to 12 degrees in one embodiment of the present invention;
FIG. 12 is a schematic representation of the results of an aircraft stall-delaying experiment at a high angle of attack on an aircraft wing in one embodiment of the present invention.
Description of the embodiments
The invention provides a multi-input multi-output non-stationary random vibration test system and a test algorithm, which are described in detail below with reference to the accompanying drawings. In the description of the present invention, it should be understood that the directions or positional relationships indicated by the terms "left", "right", "upper", "lower", "bottom", etc., are based on the directions or positional relationships shown in the drawings, are merely for convenience of describing the present invention and simplifying the description, and are not intended to indicate or imply that the apparatus or element to be referred to must have a specific orientation, be constructed and operated in a specific orientation, and "first", "second", etc. do not represent the importance of the parts and thus are not to be construed as limiting the present invention. The specific dimensions adopted in the present embodiment are only for illustrating the technical solution, and do not limit the protection scope of the present invention.
The embodiment provides an aircraft with integrated fluid thrust vectoring nozzle and wing, as shown in fig. 1, comprising an integrated flying wing section 0, an engine 1, an engine tail nozzle and the fluid thrust vectoring nozzle. The engine 1 is embedded in the flying integrated airfoil 0, and the tail nozzle of the engine is preset to be upwards deviated.
As shown in fig. 2, compared with the design that the conventional nozzle is located at the tail end of the wing, the thickness h2 of the tail of the wing is effectively reduced, so as to reduce the pressure difference resistance in the flight process.
The preset upward-offset engine tail nozzle comprises a circular-square rectifying section 21 and a square bending deflection section 22 which are sequentially connected. The rounded rectifying section 21 is intended to convert the circular jet outlet of the engine into a rectangular jet so that the tangential velocity profile of the jet matches the jet. The aspect ratio of which depends on the number of engines, the area of the jet outlet, and the flow rate of the jet outlet at the operating point.
The fluid thrust vectoring nozzle comprises a plurality of sections of nozzles, and each section of nozzle comprises a deflection wall plate, a deflection control hole, a deflection control valve and a deflection hydrostatic cavity. The number of segments depends on the thickness and deflection angle requirements of the aircraft wing aft-body. As shown in fig. 3, in this embodiment, the fluid thrust vectoring nozzle includes two sections, specifically including a first deflection wall plate 31, a first deflection control hole 33, a first deflection control valve 35, a first deflection hydrostatic chamber 37, a second deflection wall plate 32, a second deflection control hole 34, a second deflection control valve 36, and a second deflection hydrostatic chamber 38.
The wall surface of the deflection wall plate is a plane rectangle, and the length-width ratio of the wall surface is determined by the jet expansion angle and the space position of the shear layer, and is mainly obtained through an experiment mode. The sum of the areas of all deflection-control holes on the same wall surface is not more than 10% of the area of the wall surface.
As shown in fig. 3 and 4, a first deflection-control hole 33 and a second deflection-control hole 34 are provided at the front ends of the first deflection wall plate 31 and the second deflection wall plate 32, respectively; the diameter and number of deflection control holes depends on the size of the secondary flow required. The hole center distance of two adjacent deflection control holes is not more than three times of the radius of the deflection control holes. The aircraft control mode of the invention is not limited to passive fluid thrust vectors, i.e., deflection may be based on a control orifice mode or may be based on a secondary flow injection mode.
The deflection hydrostatic cavity is used for connecting the deflection control hole with the outside atmosphere. The deflection control valve is used for changing the communication degree of the deflection hydrostatic cavity.
The deflection control valve comprises a linear limit groove, a valve plate and a driving device; the linear limit groove is used for limiting the moving track of the valve plate, so that the valve plate can only move along a straight line; the driving device is used for controlling the valve plate to be at the position on the linear limit groove. The opening degree of the deflection control valve between the deflection hydrostatic cavity and the atmosphere can be changed by changing the running position of the valve plate on the linear limit groove, so that the communication degree between the deflection hydrostatic cavity and the outside atmosphere and the transported flow are adjusted, and finally the coanda degree between jet flow and the wall surface is changed, so that the jet flow direction is changed in a linear and controllable manner.
The profile of the square bending deflection section 22 is consistent with the profile of the preset upper deflection end position of the square bending deflection section 22 (namely, the middle part from the square bending deflection section 22 to a section of deflection control hole 33), and the bending angle depends on the deflection angle of the first section of the fluid thrust vectoring nozzle and the maximum deflection angle during double-section deflection. The determination method comprises the following steps: 1. determining the length of a deflection wallboard of each section of the fluid thrust vectoring nozzle and the mapping relation between the angle of the deflection wallboard and the deflection capability of jet flow through a ground wind tunnel experiment or numerical simulation; 2. comprehensively considering the maximum jet deflection angle and the single-section deflection angle during double-section deflection to determine the optimal deflection combination; 3. the thrust line during primary deflection is overlapped with the full-machine thrust line, and the outlet angle of the outlet position of the square bending deflection section 22 is equal to the jet deflection angle of primary deflection.
A method of controlling an aircraft having a fluid thrust vectoring nozzle integrated with a wing in accordance with the present invention is illustrated in figures 5 to 10. When the first and second deflection control valves 35 and 36 are opened, the engine jet will deflect upward along the square bend deflection section 22, with a preset deflection angle between the engine jet and a wall of the jet, providing lift torque and thrust due to the jet thrust line located behind the center of gravity, as shown in fig. 5 and 6. As shown in fig. 7 and 8, when the first deflection control valve 35 is closed and the second deflection control valve 36 is open, the jet will form a coanda state where the jet will be ejected along the aircraft axis and the engine thrust will be parallel to the aircraft axis providing the thrust required by the aircraft and not providing the steering torque. As shown in fig. 9 and 10, when the first-stage deflection control valve 35 and the second-stage deflection control valve 36 are both closed, the jet will double coanda and will be ejected downward along the second-stage wall surface, at which time the jet will produce a low head moment, which can provide a low head steering moment and thrust, since the jet's point of action is behind the center of gravity. Between the three states, there is a good linear relationship between deflection control valve opening and the force and torque provided, thereby providing a continuous and effective steering torque for the aircraft. The attitude control moment of the aircraft is derived from the direct force of the spray pipe on one hand and from the annular quantity control effect generated by the injection effect of the jet flow on outflow on the other hand.
As shown in fig. 11, the wind tunnel experiment result proves that the control mode described in the embodiment can provide low-head and head-up pitching control moment in the cruise attack angle of the aircraft.
As shown in fig. 12, wind tunnel experiment results prove that the control mode described in the embodiment effectively improves the stall attack angle of the wing through an inner-outer coupling mode at an attack angle of more than 16 degrees, and achieves the effect of increasing the lift force.
Stealth performance is an important index for measuring the current weaponry and is a remarkable characteristic of novel weaponry. Stealth performance includes radar stealth, infrared stealth, visible light stealth (formerly known as camouflage), laser stealth, and acoustic stealth technologies. With the development of infrared sensor technology and computer technology, the capability of the infrared detection system for detecting various targets is greatly improved, so that in the overall design process of the aircraft, infrared radiation characteristic signals must be effectively reduced according to the principle of balanced detectability while radar scattering cross sections are reduced, and the infrared stealth design of the aircraft is developed. The exhaust system is the most main infrared radiation source of the aircraft, the infrared radiation of the thermal cavity of the tail jet pipe and the tail jet flow is mainly concentrated in a medium wave band of 3-5 mu m, and the band is in the main working range of the military infrared detection system, so that the infrared stealth of the aircraft is the primary problem which must be solved. The reduction of the infrared radiation intensity of the exhaust system is an important technical approach for inhibiting the infrared radiation characteristics of the exhaust system, so the main measures of infrared stealth of the exhaust system are as follows: shielding the high-temperature component of the exhaust system, and reducing the projection area of the high-temperature component of the exhaust system in the direction of the detector; the temperature of the thermal cavity of the tail jet pipe and the temperature of the tail jet flow are reduced, and the spectral radiation degree is reduced; and the low emissivity material is adopted, so that the emissivity of the wall surface of the thermal cavity of the tail pipe is reduced. The shielding technique reduces the intensity of infrared radiation by shielding high temperature components of the exhaust system.
The first-section deflection wall plate and the second-section deflection wall plate of the embodiment shield jet flow from the lower part, the shielding effect is quite obvious, meanwhile, the temperature of the jet flow of the engine is reduced through mixing between the engine and the outflow, and the infrared stealth capability of the aircraft is enhanced.
Based on the description of the preferred embodiments of the present invention, it should be clear that the invention defined by the appended claims is not limited to the specific details set forth in the above description, but that many apparent variations of the invention are possible without departing from the spirit or scope thereof.
Claims (10)
1. An aircraft integrating a fluid thrust vectoring nozzle with a wing is characterized by comprising an aircraft integrated airfoil, an engine tail nozzle and the fluid thrust vectoring nozzle; the engine is embedded in the flying integrated wing section, and the tail nozzle of the engine is preset to be upwards deviated.
2. The integrated aircraft of claim 1, wherein the engine tail nozzle comprises a rounded rectifying section and a square bending deflection section connected in sequence.
3. The fluid thrust vectoring nozzle and wing integrated aircraft of claim 2 wherein the aspect ratio of the fairing sections is dependent on the number of engines, jet outlet area and operating point jet outlet flow size.
4. The integrated aircraft of claim 1 wherein the thrust vectoring nozzle comprises a plurality of nozzle segments, each nozzle segment comprising a deflector wall, a deflector control orifice, a deflector control valve, and a deflector hydrostatic chamber.
5. The integrated fluid thrust vectoring nozzle and wing aircraft of claim 4 wherein the wall of the deflector wall is planar rectangular and the aspect ratio of the wall is determined by the jet divergence angle and the spatial position of the shear layer and is obtained experimentally.
6. The integrated fluid thrust vectoring nozzle and wing aircraft of claim 5 wherein the sum of the areas of all deflection control apertures on the same wall is no more than 10% of the wall area.
7. The integrated fluid thrust vectoring nozzle and wing aircraft of claim 4, wherein the yaw control aperture is provided at the forward end of the yaw wall plate; the diameter and number of deflection control holes are dependent on the size of the secondary flow rate required.
8. The integrated fluid thrust vectoring nozzle and wing aircraft of claim 4 wherein the deflection hydrostatic chamber is for connecting the deflection control orifice with the ambient atmosphere.
9. The integrated fluid thrust vectoring nozzle and wing aircraft of claim 4 wherein the yaw control valve is adapted to vary the degree of communication of the yaw hydrostatic chamber.
10. The integrated aircraft of claim 9, wherein the yaw control valve comprises a linear limit slot, a valve plate, and a drive; the linear limit groove is used for limiting the moving track of the valve plate, so that the valve plate can only move along a straight line; the driving device is used for controlling the valve plate to be at the position on the linear limit groove.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311292108.8A CN117465664A (en) | 2023-10-08 | 2023-10-08 | Aircraft integrating fluid thrust vectoring nozzle and wing |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311292108.8A CN117465664A (en) | 2023-10-08 | 2023-10-08 | Aircraft integrating fluid thrust vectoring nozzle and wing |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117465664A true CN117465664A (en) | 2024-01-30 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202311292108.8A Pending CN117465664A (en) | 2023-10-08 | 2023-10-08 | Aircraft integrating fluid thrust vectoring nozzle and wing |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117465664A (en) |
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2023
- 2023-10-08 CN CN202311292108.8A patent/CN117465664A/en active Pending
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