WO2014183612A1 - 战斗机座舱透明件紧固结构及其紧固方法 - Google Patents

战斗机座舱透明件紧固结构及其紧固方法 Download PDF

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Publication number
WO2014183612A1
WO2014183612A1 PCT/CN2014/077241 CN2014077241W WO2014183612A1 WO 2014183612 A1 WO2014183612 A1 WO 2014183612A1 CN 2014077241 W CN2014077241 W CN 2014077241W WO 2014183612 A1 WO2014183612 A1 WO 2014183612A1
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WO
WIPO (PCT)
Prior art keywords
transparent member
cockpit
fastening
force arm
arm
Prior art date
Application number
PCT/CN2014/077241
Other languages
English (en)
French (fr)
Inventor
谢晓斌
李震
谢隽永
Original Assignee
Xie Xiaobin
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Xie Xiaobin filed Critical Xie Xiaobin
Publication of WO2014183612A1 publication Critical patent/WO2014183612A1/zh

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/14—Windows; Doors; Hatch covers or access panels; Surrounding frame structures; Canopies; Windscreens accessories therefor, e.g. pressure sensors, water deflectors, hinges, seals, handles, latches, windscreen wipers
    • B64C1/1476—Canopies; Windscreens or similar transparent elements
    • B64C1/1492—Structure and mounting of the transparent elements in the window or windscreen

Definitions

  • the invention relates to military aircrafts of various types of fighters in the field of national defense construction, in particular to a rigid mounting structure of a transparent part of a fighter cockpit and a fastening method thereof. Background technique
  • the fighter cockpit transparent parts generally include a fixed windshield transparent piece and a movable hatch cover transparent piece. It is not uncommon for the integrated canopy cover to be integrated into the present.
  • military aircraft has become an important manifestation of a country's military strength, and the cockpit transparent parts must have the following functions as important optical structural components on the fighter: First, it is called a structural member, and must have sufficient strength to Withstand the cockpit pressure of the fighter, the aerodynamic load, the structural load of the body and the impact load of the bird; the second is to have a good optical performance as a transparent observation window, including transmittance, haze, resolution, optical angular deviation, optical distortion, A number of important performance indicators such as ghosting, binocular parallax, and birefringence; the third is that it must have reliability of use and a long service life.
  • important functions such as bulletproof, waterproof defrost and stealth must be provided as needed.
  • the structural integrity, installation robustness and stability of use are closely related to the pilot's living environment, which will directly affect flight
  • the bolt connection is a way of opening the hole at the edge of the transparent part of the cockpit and directly connecting with the skeleton of the fuselage through hardware such as bolts.
  • the advantage is that the strength is high and the integrity is good, but the following defects are usually present:
  • the manufacturing process of the fighter aircraft has high requirements on the processing technology of the transparent parts of the cockpit. Once the processing size or curvature of the transparent part of the cockpit has a slight error, it affects the fit of the skeleton with the skeleton; or the bolt hole is opened at the edge of the transparent part of the cockpit. The center of the hole is not completely aligned with the center of the corresponding hole on the skeleton; or the bolt hole opened at the edge of the transparent part of the cockpit has defects such as cracks, notches, and poor polishing.
  • the above situation will directly lead to difficulty in installing the transparent part of the cockpit and stress after installation. Concentration, and the interchangeability of cockpit transparent products is poor;
  • the atmospheric pressure outside the cockpit decreases with the increase of the flying height. For example, the air pressure at 18,000 meters is only about one-fifth of the ground, while the cockpit maintains an atmospheric pressure to meet the physiological needs of the pilot.
  • the temperature outside the cockpit is roughly Between -60 ° C ⁇ 10 (TC change, the cabin is maintained at about 20 ° C.
  • the cockpit transparent parts under such a sharp pressure difference and temperature difference, plus the aerodynamic load, vibration and fatigue load Under the action, any small residual stress or assembly stress near the screw hole can be easily expanded into cracks or even cracks, resulting in flight accidents such as cockpit transparent parts blasting;
  • the impact load of the bird can be said to be one of the most serious external load types of the cockpit transparent part, because the overall rigidity of the cockpit transparent part and the skeleton ⁇ bolted joint is large, so the impact load of the bird collision moment is small except that the cockpit is transparent.
  • Most of the parts will be transmitted by the cockpit transparent parts to the fuselage frame through the bolts. Therefore, the stress near the bolt hole of the cockpit transparent part will increase instantaneously. The superposition of the residual stress and assembly stress will also cause cracks in the transparent part of the cabin. In severe cases, it may cause flight accidents such as damage to the cockpit transparent parts.
  • test results of the performance of the above-mentioned cockpit transparent parts show that most of the damaged parts of the transparent parts are the edge bolt holes, and the test components specified in the standard are compared with the transparent parts of the cockpit used on the fighter aircraft. Assembly stress, the test environment is also far less complex than the actual flight environment. It can be seen that the performance of the cockpit transparent part in the actual flight state is not up to the value in the above table, and it is uncertain and difficult to detect, thus opening the hole in the cockpit transparent part and passing the bolt and the fuselage The way the skeleton is connected, to the cockpit transparent part itself and even the entire cockpit stress structure The security breach is quite obvious.
  • inorganic glass itself has good wear resistance, heat resistance and corrosion resistance. It is usually used for parts with high aerodynamic heat and wear resistance. However, inorganic glass is brittle. The material is rarely used throughout the cockpit cover of the fighter; it is difficult to punch holes in inorganic glass and sensitive to stress concentration at the hole, which also limits its use in the cockpit transparent.
  • the fastening structure of the cockpit transparent parts in the world is almost the same, but the material and manufacturing process of the cockpit transparent parts are improved.
  • the fastening structure of the cockpit transparent parts and the fuselage skeleton has not changed.
  • the above problems still exist, the fighters
  • the cockpit transparent fastening structure has also become a major problem for the related art.
  • the object of the present invention is to overcome the shortcomings of the prior art that the punching of the transparent part of the fighter cockpit is easy to cause stress concentration, and to provide a new non-perforated cockpit transparent member fastening structure.
  • the present invention discloses a fighter cockpit transparent member fastening structure, which comprises a body and a cockpit transparent member mounted on the body; the cockpit transparent member has a frame at the periphery thereof, and the frame is formed on the frame In conjunction with the mounting structure of the body, the frame includes a pressing assembly and a fastening assembly, and the fastening assembly generates a pre-stress by the cooperation of the pressing assembly and the cabin transparent member to thereby fasten the cabin transparent member.
  • the fastening assembly includes two arcuate arms symmetrically clamped to the cabin transparent member, and the two arcuate arms are interposed to form a surrounding space
  • the arcuate arm includes a first arm a second force arm connected to the first force arm, a joint of the first force arm and the second force arm forming a slip end, the first force arm being away from the second force arm Forming a compression end on one side, the second force arm forming a fastening end on a side away from the first force arm, the pressure end of the first force arm receiving the compression of the compression assembly and cooperating with the
  • the cockpit transparency drives the first and second force arms to generate a pre-stress.
  • a further improvement of the present invention is that the pressing assembly comprises a first pressure strip and a second pressure strip;
  • the first pressure bar is disposed on an outer side of the first force arm of the arcuate arm; the two sliding ends of the arcuate arm of the fastening component abut against the first pressure bar;
  • the pressing end abuts against the second pressure strip, and the two fastening ends of the arcuate arm abut against the two sides of the connecting portion of the cabin transparent member;
  • the first pressure bar and the second pressure bar respectively open a plurality of corresponding bolt holes; the first pressure bar and the second pressure bar are fastened by bolts; the second pressure bar presses the The two compression ends of the arcuate arm are displaced toward the first pressure bar, and the two sliding ends of the arcuate arm are displaced away from each other, and the two fastening ends of the arcuate arm are restricted by the transparent part of the cabin Thereby driving the first force arm and the second force arm to generate pre-stressed fastening of the cabin transparency.
  • a further improvement of the present invention is that the first pressure bar forms a slot-shaped hole and is fastened to the second pressure bar by a bolt penetrating in the slot-shaped hole, and the second pressure bar and the Positioning the fastening component in a first direction through the slotted hole;
  • the cabin transparent member performs a second direction and a third direction position adjustment through the enclosure space.
  • the cabin transparency comprises an inner layer and an outer layer bonded to both sides of the inner layer; the outer layer is recessed to form the connecting portion.
  • a further improvement of the present invention is that the cabin transparency comprises an inner layer and an outer layer bonded to both sides of the inner layer; the inner layer extends outward to form the connection protruding from the outer layer unit.
  • a further improvement of the present invention is that the surface of the cabin transparent member forms the connecting portion that protrudes outward.
  • the first force arm is a short straight arm and the second force arm is an arcuate arm.
  • a pressing plate is coupled to the fastening end, and a connecting area of the pressing plate and the second force arm is recessed inward to form a platen position adjusting area.
  • a further improvement of the present invention is that a glue (such as a UV glue) or a double-sided tape (such as a 3M glue) or a cushion (such as a UV glue) can be applied between the two fastening ends and the transparent part of the cabin (such as Rubber sheet).
  • a glue such as a UV glue
  • a double-sided tape such as a 3M glue
  • a cushion such as a UV glue
  • the second force arm is spaced apart to form a plurality of overflow grooves; the enclosed space is filled with a sealant.
  • the sliding end of the arcuate arm has a circular arc surface or a sloped surface.
  • a further improvement of the invention is that the thickness of the second force arm forms a thick to thin gradient from the slip end to the fastening end.
  • the compressed end of the arcuate arm of the fastening assembly extends to form a rotational positioning rib, and the second pressure strip is formed with a rotary positioning groove corresponding to the rotational positioning edge of the fastening component.
  • a further improvement of the invention consists in that the pressed ends of the arcuate arms are connected by an arcuate deformation zone.
  • the transparent part of the cockpit is fastened by the frame and the frame of the fuselage, and the bolt hole is no longer needed on the transparent part of the cockpit, and the internal stress balance of the transparent part of the cockpit is not damaged, and the original strength of the transparent part of the cockpit is maintained.
  • There is no weak point of stress no assembly difficulties due to machining defects of bolt holes, and stress concentration during assembly; there is no superposition and expansion of the above stress during use.
  • the cockpit transparent member is fastened to the fuselage frame in a positionally adjustable manner, and the fastening assembly has a certain tolerance for the manufacturing error of the cockpit transparent member, thereby assembling the cockpit transparent member and the skeleton More convenient and avoid assembly
  • the generation of stress can greatly improve the assembly interchangeability of the transparent parts of the cabin.
  • the invention bites the plane of the transparent part of the cockpit through the prestressed structure, and generates a rigid frame integrated with the transparent part of the cockpit, and fixes the transparent part of the cockpit to the fuselage frame through the frame, thereby forming a
  • the new integrated force structure of the cockpit transparent part with pre-stress buffer function and the cockpit effectively transmits the load on the transparent part of the cockpit to the fuselage through the prestressed structure, ensuring the integrity of the cockpit structure.
  • the edge of the transparent part of the cockpit of some fighter aircraft has a circular arc surface.
  • the second force arm of the fastening component is split into a plurality of jaws, so that the fastening end of the fastening component can be more closely attached to the arc surface of the transparent part of the cabin, so that the fastening component does not damage the cabin transparent
  • the pre-stress is generated more firmly and stably to fasten the transparent part of the cabin.
  • the transparent part of the cockpit of the invention is fastened by the prestressed structure and the skeleton of the fuselage, and the fastening component is selected from materials having considerable strength and elasticity and toughness, and the cockpit transparent part and the fuselage frame are thermally expanded during flight.
  • the prestressing compensation function of the fastening component itself can be buffered, the destructive force generated by the deformation difference between the two is effectively solved, and the safety and stability of the cabin structure are maintained.
  • the fastening component is selected from materials having considerable strength and at the same time having certain elasticity and toughness, and during the flight, the deformation and stress of the cockpit transparent member are subjected to loads such as pressure difference, temperature difference, and bird impact. It can be buffered by the release and re-generation process of the pre-stress contained in the fastening component itself, which not only does not superimpose various complicated stresses on each other, but can eliminate or reduce the stress concentration to a certain extent. Keep the cockpit structure safe and stable.
  • the present invention has the above-mentioned characteristics, so that the problem of difficulty in installing complex cockpit transparent parts such as integral molding can be solved correspondingly, and at the same time, the stability of the cockpit transparent member is improved, the life is extended, and the maintenance strength is lowered. Made a positive contribution.
  • the prestressed fastening process of the present invention is to pre-stress the fastening component by tightening the relevant bolts.
  • the selection of the raw materials of each component and the geometric design are adopted in the previous design module. After the workers can tighten the relevant bolts in place, the preset tightening force can be obtained without being affected by uncertain factors such as the operating force, which greatly reduces the operating conditions and technical requirements.
  • FIG. 1 is a schematic view showing the overall structure of a fighting machine when a whole block cockpit transparent member structure is used in the first embodiment of the fighter cockpit transparent member fastening structure according to the present invention
  • Figure 2 is a front elevational view of the fighter cockpit of Figure 1;
  • Figure 3 is a cross-sectional view of Figure 2;
  • Figure 4 is a schematic view showing the overall structure of the fighter when the multi-seat cockpit transparent member structure of the first embodiment of the fighter cockpit transparent member fastening structure of the present invention is used;
  • Figure 5 is a front view of the fighter cockpit of Figure 4.
  • Figure 6 is a cross-sectional view of Figure 5;
  • Figure 7, Figure 23, Figure 29, Figure 35 is a partial enlarged view of various embodiments of the area A in Figure 6;
  • Figure 8 is a plurality of cockpit transparent parts for the first embodiment of the fuser cabin transparent member fastening structure of the present invention Side view of the cockpit of the structure;
  • Figure 9 is a cross-sectional view of Figure 8.
  • Figure 11, Figure 25, Figure 31, Figure 37 are partial enlarged views of various embodiments of the C region of Figure 9;
  • Figure 13 is a perspective view showing the overall connection structure of the cockpit transparent member and the frame of the cockpit transparent member fastening structure of the present invention
  • Figure 14, Figure 21, Figure 27, Figure 33 are perspective views of various embodiments of the partial connection structure of the cockpit transparent member and the frame of the fuselage hull transparent member fastening structure of the present invention
  • Figure 15 is an exploded view of Figure 14;
  • Figure 16 is a plan view showing the fastening assembly of the present invention.
  • Figure 17 is a perspective view of the fastening assembly of the present invention.
  • Figure 18 is a perspective view of a second pressure strip of the present invention.
  • Figure 19 is a schematic view showing the compression deformation of the arc deformation zone of the fastening assembly of the present invention.
  • Figure 20 is a schematic view showing the principle of the fastening process of the frame and the transparent part of the cabin in the transparent structure of the fighter cockpit of the present invention
  • Figure 22 is an exploded view of Figure 21;
  • Figure 28 is an exploded view of Figure 27;
  • Figure 34 is an exploded view of Figure 33;
  • Figure 39 is a side view of the fighter cockpit of the second embodiment of the fuselage hull transparent member fastening structure of the present invention.
  • Figure 40 is a cross-sectional view of Figure 39;
  • Figure 41 is a partial enlarged view of the C area in Figure 40. detailed description
  • a fighter cockpit transparent member fastening structure of the present invention includes a body 1 and a cabin transparent member 2 mounted on the body 1; a cockpit transparent member 2 The periphery is combined with a frame 3.
  • the frame 3 is formed with a mounting structure for engaging the body 1.
  • the frame 3 includes a pressing assembly 31 and a fastening assembly 32.
  • the pressing member 31 and the cabin transparent member 2 cooperate to compress the fastening assembly 32 to generate a pre-stress and thereby fasten the cabin transparent member 2 .
  • the combat machine can use the structure of the whole cockpit transparent member 2 as shown in Fig. 1; the structure of the plurality of cockpit transparent members 2 as shown in Fig. 8 can also be used.
  • the horizontal mounting direction of the cockpit transparent member 2 in Figures 14 and 15 is taken as the X-axis direction, and the thickness direction of the cockpit transparent member 2 is taken as the Y-axis direction.
  • the vertical mounting direction of the cockpit transparent member 2 is taken as the Z-axis direction, and the X-axis is perpendicular to the Y-axis, and the Z-axis is perpendicular to the plane formed by the X-axis and the Y-axis;
  • the fastening component 32 includes two arcuate arms 321 symmetrically clamped to the cabin transparent member 2, and the material thereof should be selected from materials having considerable strength and elasticity and toughness, such as metal, engineering plastics, polymer materials, etc.; Bow arm 321
  • the enclosing space 320 is formed between the first force arm 3211 and the second force arm 3212 connecting the first force arm 3211.
  • the connection between the first force arm 3211 and the second force arm 3212 forms a
  • the sliding end 3213, the sliding end 3213 has a circular arc surface or a sloped surface to ensure less resistance during the sliding process; the first force arm 3211 forms a pressure receiving end 3214 on a side away from the second force arm 3212.
  • the pressing end 3214 is extended to form a rotating positioning edge 3217; the second force arm 3212 forms a fastening end 3215 on a side away from the first force arm 3211, the pressing end 3215 is coupled with a pressing plate 3216, and the pressing plate 3216 and the second
  • the connecting portion of the arm 3212 is recessed inwardly to form a platen position adjusting portion 3218.
  • the platen position adjusting portion 3218 can realize the slight self-position adjustment of the platen 3216 during the fastening process, so that the flat seat transparent member is attached more smoothly.
  • the pressed end 3214 of the first force arm 3211 receives the compression of the compression assembly 31 and cooperates with the cabin transparent member 2 to drive the first force arm 3211 and the second force arm 3212 to generate a pre-stress.
  • the first force arm 3211 is a short straight arm
  • the second force arm 3212 is an arcuate arm
  • the thickness of the second force arm 3212 forms a thickness from the sliding end 3213 to the fastening end 3215. With a thin gradient, this structure ensures that the entire curved arm is fully and evenly deformed and is not easily broken.
  • the two arcuate arms 321 are connected between the two pressure receiving ends 3214 by providing an arc deformation region 3219.
  • the arc deformation region 3219 is pressed from the arc shape. Extending, the deformation process of the curved deformation zone 3219 is shown in FIG. 19; the design of the arc deformation zone 3219 ensures that the fastening component 32 has a certain extension space; the two compression ends 3214 of the fastening component 32 cooperate A plurality of bolt holes are formed. Between the two fastening ends 3215 and the cabin transparent member 2, an adhesive (such as UV glue) or a double-sided adhesive (such as 3M glue) or a cushion (such as a rubber sheet) may be applied. The second force arm 3212 is spaced apart to form a plurality of overflow grooves 32121.
  • an adhesive such as UV glue
  • a double-sided adhesive such as 3M glue
  • a cushion such as a rubber sheet
  • overflow groove 32121 serves to overflow the excess unsolidified sealant when the fastening component 32 is clamped and prevent the expansion or contraction of the sealant during the solidification process.
  • the effect of the pre-stress generated by the fastening assembly 32 ensures pre-stressed fastening between the fastening assembly 32 and the cabin transparency 2 and at the same time achieves a seal.
  • the pressing assembly 31 includes a first pressure strip 311 and a second pressure strip 312; the surface of the second pressure strip 312 cooperates with the rotating positioning edge 3217 to provide two long rotating positioning grooves.
  • the radius of the rotary positioning groove 3121 is equal to or slightly larger than the radius of the rotary positioning edge 3217, so that when the entire transparent member fastening structure is in the pre-fastening and fastening state, the rotary positioning edge 3217 can effectively rotate the positioning groove.
  • the positioning and rotation in the 3121, the two sliding ends 3213 will only be displaced in the thickness direction of the cabin transparent member 2 on the surface of the first pressure bar.
  • the first pressure bar 311 is disposed on the outer side of the first force arm 3211 of the arcuate arm 321; the two sliding ends 3213 of the arcuate arm 321 of the fastening component 32 abut against the first pressure bar 311; the two pressure ends of the arcuate arm 321 3214 abuts against the second pressure strip 312, and the two fastening ends 3215 of the arcuate arm 321 abut against both sides of the joint of the cabin transparent member 2.
  • the first pressure bar 311 and the second pressure bar 312 respectively open a plurality of corresponding bolt holes; the first pressure bar 311 and the second pressure bar 312 are fastened by bolts; the second pressure bar 312 presses the two pressures of the bow arm 321
  • the end 3214 is displaced toward the first pressure bar 311, and the two sliding ends 3213 of the arcuate arm 321 are displaced from each other.
  • the two fastening ends 3215 of the arcuate arm 321 are restricted by the cabin transparent member 2, thereby driving the first force arm.
  • the 3211 and the second force arm 3212 generate a pre-stressed fastening cabin transparency 2.
  • the first pressure strip 311 forms a slotted hole 3111 and is fastened to the second pressure strip 312 by a bolt passing through the slotted hole 3111, and the second pressure strip 312 and the fastening component 32 are passed along the slotted hole 3111.
  • the position adjustment is performed in the Y-axis direction; the cabin transparent member 2 is adjusted in the X-axis direction and the Z-axis direction by the enclosed space 320.
  • the first pressure bar 311 forms a pressing end 3112, and the two sides of the pressing end 3112 form a connecting end 3113.
  • the connecting end 3113 is a mounting structure for the body 1, and the first pressure bar 311 is fixed to the body 1 through the connecting end 3113.
  • the installation structure can also use other Connection structure.
  • a sealant is filled in the gap between the second force arm 3212, the cabin transparent member 2, and the second pressure bar 312, thereby achieving more stable fastening; since the overflow groove 32121 is opened on the second force arm 3212.
  • the use of the overflow groove 32121 serves to overflow the excess unsolidified sealant when the fastening assembly 32 is clamped and to prevent the expansion or contraction of the sealant during the solidification process from affecting the pre-stress generated by the fastening assembly 32.
  • the second pressure strip 312 is placed in the enclosed space 320 of the fastening assembly 32, and the rotary positioning groove 3121 is engaged with the rotary positioning edge 3217, and then The sealing space 320 is filled with a sealant and the first force arm 3211 of the fastening component 32 is disposed on the first pressure bar 311.
  • the two fastening ends 3215 of the arcuate arm 321 abut against the two sides of the connecting portion of the cabin transparent member 2
  • the first pressure bar 311 is disposed on the outer side of the first force arm 3211; the two sliding ends 3213 of the arcuate arm 321 abut against the first pressure bar 311, and the two pressure receiving ends 3214 of the arcuate arm 321 abut against the second
  • the outer side surface of the pressure strip 312 is pre-tensioned by bolts sequentially passing through the bolt holes of the first pressure strip, the fastening assembly 32 and the second pressure strip 311, and the second pressure strip 312 and the fastening assembly 32 pass through the slotted hole 3111.
  • the position adjustment is performed along the Y-axis direction; the cabin transparent member 2 is adjusted in the X-axis direction and the Z-axis direction by the enclosure space 320. After the position of the cabin transparent member 2 is adjusted into position, the first pressure strip 311 is fastened by the bolt. And the second pressure bar 31 2 to complete the fastening.
  • the two pressure receiving ends 3214 of the arcuate arms 321 are displaced in the direction of the first pressure bar 311 by the compression of the second pressure bar 312, by rotating the positioning edges 3217 and rotating.
  • the engagement of the positioning groove 3121 ensures that the pressure receiving end 3214 is only displaced in the Z-axis direction during the movement, and the distance between the pressure receiving ends 3214 of the two arcuate arms 321 is controllable (unchanged) during the fastening process.
  • the two sliding ends 3213 are displaced away from each other by the inner side surface of the first pressure bar 311, and the two fastening ends 3215 are displaced adjacent to each other until they abut against the side of the cabin transparent member 2, so the two fastening ends
  • the distance between the platens 3216 of the 3215 is also controllable, and the fastening point on the cockpit transparent member 2 is also controllable; further pressing the two pressure receiving ends 3214 to the direction of the first pressure bar 311, thereby driving the two slips
  • the ends 3213 continue to move away from each other, and the two fastening ends 3215 abut against the side of the cabin transparent member 2 and are thus restrained, and the first force arm 3211 and the second force arm 3212 are thereby deformed and To prestress, thus having a transparent cockpit stable prestressed structure member 2 and the frame 3 reaches fastened state, the transparent cockpit fastening member 2 is obtained.
  • the cabin transparent member 2 is a sandwich structure including an inner layer 21 and an outer layer 22 bonded to both sides of the inner layer; the outer layer 22 forms a recess and the outer layer 22 is flush with the inner layer 21, tightly
  • the solid end 3215 is attached to the recessed portion, and the recessed portion is the connecting portion of the cabin transparent member 2, that is, the outer layer is recessed to form a connecting portion.
  • the outer layer 22 may also not form a recessed portion, and the fastening end 3215 is directly in close contact with the outer surface of the outer layer 22.
  • the connecting portion is the edge of the cockpit transparent member 2, and the connection structure of the cockpit transparent member 2 and the frame 3 is shown in FIG. 21- 26.
  • the inner layer 21 extends outwardly to form a joint protruding from the outer layer 22; the fastening end 3215 closely fits to both sides of the inner layer 21 protruding outer layer 22; the connection structure of the cabin transparent member 2 and the frame 3 See Figure 27-32.
  • a fighter cockpit transparent member fastening structure of the present invention is a single cockpit transparent member 2, and the main structure thereof is the same as that of the first embodiment.
  • the inner side of the cockpit transparent member 2 is further laminated with a plurality of reinforcing ribs 23 arranged along the curved line of the cabin transparent member 2 in the same plane; the first end of the reinforcing rib 23 is attached to the inner surface of the cockpit transparent member 2; The second ends of the ribs 23 form the ribs 231 and are fastened to each other by the bezel 3.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Connection Of Plates (AREA)
  • Tents Or Canopies (AREA)
  • Window Of Vehicle (AREA)

Abstract

一种战斗机座舱透明件紧固结构及其紧固方法，紧固结构包括机体（1）和安装于机体（1）上的座舱透明件（2），座舱透明件（2）周边结合有一边框（3），边框（3）包括压迫组件（31）和紧固组件（32），通过压迫组件（31）和座舱透明件（2）的配合压迫紧固组件（32）生成预应力进而紧固座舱透明件（2）。该紧固结构强度高，适应面广。

Description

战斗机座舱透明件紧固结构及其紧固方法 技术领域
本发明涉及国防建设领域中的各类战斗机等军用飞机，尤指一种战斗机座舱透明件的紧 固安装结构及其紧固方法。 背景技术
战斗机座舱透明件一般包括固定风挡透明件及活动舱盖透明件， 目前将两者合为一体的 整体式座舱盖也已不鲜见。 在当前形势下， 军用飞机已成为一个国家军事实力的重要体现， 而座舱透明件作为战斗机上重要的光学结构件必须具备以下多种功能： 首先既称为结构件， 必须具有足够的强度， 以承受战斗机座舱压力、 气动载荷、 机体结构载荷以及飞鸟撞击荷载 等； 第二是作为透明观察窗， 必须具有良好的光学性能， 包括透光度、 雾度、 分辨率、 光学 角偏差、 光学畸变、 重影、 双目视差、 双折射等多项重要的性能指标； 第三是必须具有使用 可靠性和较长的使用寿命。 此外， 根据需要还须具有防弹、 防水去霜及隐身等重要功能， 其 结构的完整性、 安装的牢固性和使用的稳定性与飞行员的生存环境密切相关， 将直接影响到 飞行安全和军事任务的完成。
随着科技的发展，作为座舱透明件原材料的航空玻璃，其自身技术已经有了长足的进步， 而关于座舱透明件与机身骨架的紧固安装结构或者说边缘连接方式， 目前仍主要釆用以下几 种形式： 螺栓连接、 软连接、 纯气密连接、 膨胀接头、 U型槽连接和金属插入件加强等等， 或单独使用、 或组合使用。
其中的螺栓连接即在座舱透明件边缘开孔并通过螺栓等五金件直接与机身骨架硬连接 的方式， 优点是强度较高、 整体性较好， 但是通常存在以下缺陷：
当战斗机处于地面静止状态时：
1、 战斗机生产制造过程对座舱透明件的加工工艺要求较高， 一旦座舱透明件的加工尺 寸或弧度等出现少许误差而影响了其与骨架的贴合度；或者在座舱透明件边缘开设螺栓孔的 中心与骨架上相应孔的中心没有完全对准； 或者在座舱透明件边缘开设的螺栓孔存在裂紋、 缺口、抛光不好等缺陷， 以上情况都会直接导致座舱透明件安装困难、安装后存在应力集中， 且座舱透明件制品的互换性差；
2、 座舱透明件在生产制造及加工过程中或多或少存在一定的缺陷或残余应力， 在运输、 仓储或装配的过程都可能导致该缺陷或应力逐渐放大，若没有及时发现并进行有效处理或报 废， 一旦座舱透明件产生裂紋会使其抗拉强度和抗冲击韧性等性能明显下降， 给将来的使用 带来极大的安全隐患；
3、 在座舱透明件边缘开孔就已经破坏了透明件起初的内应力平衡， 螺孔附近成为强度 上的薄弱环节， 再加上可能叠加的加工缺陷与装配应力， 透明件上很容易产生裂紋， 不但存 在隐患而且对维修保养工作也是很严峻的考验。
当飞机处于空中飞行状态时：
1、 座舱外大气压力随飞行高度升高而减小， 比如在 18000米高空的气压大约只有地面 的十五分之一， 而座舱内为满足飞行员生理需要始终保持一个大气压左右； 座舱外温度大致 在 -60°C~10(TC之间变化， 座舱内则基本维持在 20°C左右。 座舱透明件在如此剧烈的压差及 温差条件下， 加上气动荷载、 振动以及疲劳荷载等的共同作用下， 上述螺孔附近的任何微小 残余应力或装配应力，都很容易被扩展为裂紋，甚至裂缝而导致座舱透明件爆破等飞行事故；
2、 因为座舱透明件与 1机身金属骨架的热膨胀系数是不同的， 而且随温度变化的规律也 存在很大差异， 所以在空中压差与温差急剧变化时， 除了刚性较大的骨架会限制座舱透明件 的横向变形， 导致座舱透明件与骨架连接处的侧向应力增大外， 两者的接触面附近还会出现 其他复杂的应力状态， 时间长了就容易在螺孔等应力薄弱处产生疲劳裂紋， 直到座舱透明件 破坏导致灾难性的后果； 另外层合结构座舱透明件中不同材料的热膨胀系数也不尽相同， 复 杂条件下内外层的变形也不一样， 同样会对座舱透明件尤其是螺孔附近的应力状态造成影响 甚至破坏；
3、 由于螺栓会传热， 战斗机若长时间暴露于高温下， 温度会从机身通过螺栓和衬套传 到螺栓孔表面而使孔的有效支承面积变小，加之战斗机使用过程中座舱透明件上螺栓孔的不 断磨损与变形， 都对整个结构的强度与稳定非常不利。
4、 飞鸟撞击荷载可以说是座舱透明件最严重的外荷载类型之一， 因为座舱透明件与骨 架釆用螺栓连接处的整体刚性较大， 故鸟撞瞬间的冲击荷载除小部分由座舱透明件吸收外， 大部分将由座舱透明件通过螺栓传递至机身骨架， 因此座舱透明件螺栓孔附近的应力将瞬间 增大， 与上述残余应力、 装配应力等叠加后同样会使座舱透明件产生裂紋， 严重时可造成座 舱透明件破坏等飞行事故。
另外， 以下为我国航空工业总公司于一九九六年发布并实施的航空工业标准《飞机座舱 透明件设计手册》 HB/Z 290-96第 6.1.3单元， 各种典型边缘连接件性能汇总： 孔边距 12mm典型结构元件常温抗拉强度
丙 烯 酸 胶
kg/7.2cm 破 坏 部 位
1 2180 2150 螺 栓 孔 破 坏
2 1870 2240 螺 栓 孔 破 坏
3 1910 2340 螺 栓 孔 破 坏
4 2520 2000 螺 栓 孔 破 坏
5 2120 2300 螺 栓 孔 破 坏 平 均 值 2120 2246 孔边距 15mm元件原始状态常温抗拉强度
1 丙 - 1 胶 1 丙 烯 酸 胶
kg/7.2cm ！ kg/7.2cm 破 坏 部 位 螺 栓 孔 破 坏 螺 栓 孔 破 坏
3 2670 2300 螺 栓 孔 破 坏
4 2730 2330 螺 栓 孔 破 坏
5 1 238 o0 1 2830 螺 栓 孔 破 坏 平 均 值 ： 2644 2560
Figure imgf000005_0001
孔边距 12mm典型结构元件经 80°C
热处理后室温疲劳寿命
丙 - 1 胶 丙 烯 酸 胶
破 坏 部 位 破 坏 部 位 (次） (次)
207 单孔边缘处 3320 边 缘 孔 处
236 单孔边缘处 558 边 缘 孔 处
347 单孔边缘处 738 边 缘 孔 处
378 单孔边缘处 1778 边 缘 孔 处
692 单孔边缘处 841 边 缘 ？ L 丝
548 边 缘 孔 处
854 边 缘 孔 处
3220 边 缘 孔 处
503 边 缘 孔 处
1104 边 缘 孔 处
表 5 孔边距 15mm典型结构元件原始状态室温疲劳寿命
丙 - 1 胶 丙 烯 酸 胶
破 坏 部 位 破 坏 部 位 (次） (次）
11440 加强片过渡处 9454 边 缘 孔 处
9614 10368 边 缘 ？ L 丝
1782 2552 边 缘 孔 处
1745 未 断 停 载 1984 边 缘 孔 处
1451 未 断 停 载 3906 边 缘 孔 处
1952 边 缘 孔 处
4166 边 缘 孔 处 表 6 孔边距 15mm典型结构元件经 80°C χ 6h
热处理后室温疲劳寿命
丙 - 1 胶 丙 烯 酸 胶
破 坏 部 位 破 坏 部 位 (次） (次）
360 单孔边缘 11300 未断停载
560 单孔边缘 12000 未断停载
550 单孔边缘 10000 未断停载
10530 未断停载
7268 未断
15211 未断停载
14718 未断停载 表 7 孔边距 12mm典型元件湿热老化后室温疲劳寿命 丙 烯 酸 胶 （次) 破 坏 部 位
1 13277 孔 边 缘 处
2 10730 孔 边 缘 处
3 14394 孔 边 缘 处
4 6976 孔 边 缘 处
5 19867 孔 边 缘 处 表 8 典型结构元件原始状态及单面受热，单面受冷的情况下
Figure imgf000007_0001
上述座舱透明件各项性能的测试结果显示， 透明件的破坏部位绝大多数为边缘螺栓孔 处， 而且该标准规定的试验元件与运用于战斗机上的座舱透明件相比， 结构筒单、 没有装配 应力， 试验的环境也远不如实际飞行环境那么复杂。 可见， 处于实际飞行状态中的座舱透明 件的各项性能是达不到上述表格中的数值的， 而且是不确定的和难以检测的， 因而在座舱透 明件上开孔并通过螺栓与机身骨架连接的方式，对座舱透明件本身乃至整个座舱受力结构的 安全性破坏作用是相当明显的。 由此可见， 现有透明件通过边缘打孔来安装的技术， 是制约 透明件实现其安全性能的根本问题； 反则， 以无孔形式的牢固安装， 是成就透明件实现其安 全性能的关键技术贡献。
还有关于座舱透明件材料的选择， 无机玻璃本身具有良好的耐磨性、 耐热性和抗腐蚀性 等， 通常用于气动热及耐磨损要求较高的部位， 然而由于无机玻璃属于脆性材料， 很少在整 个战斗机座舱盖中全部使用；而且在无机玻璃上打孔安装较困难、对打孔处应力集中较敏感， 也都限制了其在座舱透明件中的使用。
而航空有机玻璃虽然近年来发展较快， 性能也有较大提升， 但仍难从根本上改变其硬度 低、 耐磨性差、 热膨胀系数大、 导热性差、 抗老化性能差、 防静电性能差等缺点， 其在飞机 座舱透明件中的应用也受到了一定的限制，某些特殊机型或是关键部位仍无法完全取代无机 玻璃。
近年来一些先进的战斗机越来越多的釆用一体成型全景式流线型座舱盖玻璃，其优点是 飞行员的视野更开阔、 隐身效果更好， 而且更适合快速反应与超音速飞行， 当然随之而来的 缺陷是制造难度大、 安装难度大以及有待改进的光学畸变与使用寿命问题等等， 另外一体成 型的舱盖玻璃一旦存在内应力或细小的裂紋， 将比多框架结构的玻璃更容易扩散、 维护更困 难。
目前世界上战斗机座舱透明件的紧固结构大致相同，只是在座舱透明件的材质及制造工 艺上有所改进， 座舱透明件与机身骨架的紧固结构并无改变， 上述问题依然存在， 战斗机座 舱透明件紧固结构亦成为困扰相关技术人员的一大难题。
然而随着时代的进步， 军用飞机在国家军事实力中的地位越来越重要， 如果上述问题没 有合理的解决方案， 那么飞机的发展与进步也将受到影响。 针对此类关系到国家军事实力的 重要问题， 目前尚无比较合理的解决方式， 而本发明填补了此领域的空白。 发明内容
本发明的目的在于克服现有技术在战斗机座舱透明件边缘打孔进行紧固安装容易产生 应力集中的缺陷， 而提供一种全新的不打孔的座舱透明件紧固安装结构。
为解决上述技术问题， 本发明公开了一种战斗机座舱透明件紧固结构， 包括机体以及安 装于所述机体上的座舱透明件； 所述座舱透明件周边结合有一边框， 所述边框上形成有配合 所述机体的安装结构， 所述边框包括压迫组件和紧固组件， 通过所述压迫组件和座舱透明件 的配合压迫所述紧固组件生成预应力进而紧固所述座舱透明件。
本发明的进一步改进在于， 所述紧固组件包括两个对称夹持于所述座舱透明件的弓形 臂， 两弓形臂之间夹设形成一围合空间， 所述弓形臂包括第一力臂与连接所述第一力臂的第 二力臂， 所述第一力臂与所述第二力臂的连接处形成滑移端， 所述第一力臂于远离所述第二 力臂的一侧形成受压端， 所述第二力臂于远离所述第一力臂的一侧形成紧固端， 所述第一力 臂的受压端接受所述压迫组件的压迫并配合所述座舱透明件驱使所述第一力臂与第二力臂 生成预应力。
本发明的进一步改进在于， 所述压迫组件包括一第一压力条和一第二压力条； 所述第一压力条设置于所述弓形臂的第一力臂的外侧；所述紧固组件的弓形臂的两滑移 端抵靠于所述第一压力条； 所述弓形臂的两受压端抵靠于所述第二压力条， 所述弓形臂的两 紧固端抵靠于所述座舱透明件连接部的两侧面；
所述第一压力条和所述第二压力条分别开设有复数个对应的螺栓孔；通过螺栓紧固所述 第一压力条与所述第二压力条；所述第二压力条压迫所述弓形臂的两受压端向所述第一压力 条方向位移， 所述弓形臂的两滑移端发生相互远离的位移， 所述弓形臂的两紧固端受到所述 座舱透明件的限位， 从而驱使所述第一力臂与所述第二力臂生成预应力紧固所述座舱透明 件。
本发明的进一步改进在于， 所述第一压力条形成槽形孔， 并通过穿设于所述槽形孔内的 螺栓与所述第二压力条紧固，且所述第二压力条和所述紧固组件通过所述槽形孔沿一第一方 向进行位置调整；
所述座舱透明件通过所述围合空间进行一第二方向与一第三方向的位置调整。
本发明的进一步改进在于， 所述座舱透明件包括一内层和结合于所述内层两侧的外层； 所述外层凹陷形成所述连接部。
本发明的进一步改进在于， 所述座舱透明件包括一内层和结合于所述内层两侧的外层； 所述内层向外延伸形成自所述外层之间凸出的所述连接部。
本发明的进一步改进在于， 所述座舱透明件的表面形成向外凸出的所述连接部。
本发明的进一步改进在于， 所述第一力臂为一短直臂， 所述第二力臂为一弧形臂。 本发明的进一步改进在于， 所述紧固端上结合有压板， 且所述压板与所述第二力臂的连 接区域向内凹陷形成一压板位置调节区。
本发明的进一步改进在于， 所述两紧固端与所述座舱透明件之间可涂抹粘结胶（如 UV 胶）或夹设双面胶（如 3M胶）或塾设緩冲垫（如橡胶片） 。
本发明的进一步改进在于， 所述第二力臂间隔形成复数个溢流槽； 所述围合空间内填充 有密封胶。
本发明的进一步改进在于， 所述弓形臂的滑移端呈圆弧面或斜面。
本发明的进一步改进在于，所述第二力臂的厚度自所述滑移端至所述紧固端形成一由厚 至薄的渐变。
本发明的进一步改进在于， 所述紧固组件的弓形臂的受压端延伸形成有一旋转定位棱， 所述第二压力条对应所述紧固组件的所述旋转定位棱形成有旋转定位槽。
本发明的进一步改进在于， 所述弓形臂的受压端之间通过一弧形变形区连接。
本发明由于釆用了以上技术方案， 使其具有的有益效果是：
1. 本发明中座舱透明件通过边框与机身骨架实施紧固， 座舱透明件上不再需要开设螺 栓孔， 不会破坏座舱透明件自身的内应力平衡， 保持了座舱透明件原有的强度、 没有应力薄 弱点， 不会产生因为螺栓孔的加工缺陷等而导致的装配困难以及装配过程的应力集中； 在使 用过程中也不存在上述应力的叠加与扩大。
2. 在本发明中， 座舱透明件以位置可调的方式紧固于机身骨架， 且紧固组件对座舱透 明件的生产制造误差有一定的宽容度， 因此使得座舱透明件与骨架的装配更便捷、避免装配 应力的产生， 同时可大幅提高座舱透明件制品的装配互换性。
3. 本发明通过预应力结构咬合座舱透明件平面， 并在座舱透明件周边生成了一圏与其 合为一体的刚性边框， 并通过该边框将座舱透明件固定于机身骨架， 由此形成一种全新的具 备预应力緩冲功能的座舱透明件与座舱一起的整体受力结构，有效地将座舱透明件承受的荷 载通过预应力结构传递至机身， 保证了座舱结构的完整性。
4. 在实际运用当中， 部分战斗机座舱透明件边缘呈圆弧面。 在本发明中， 紧固组件第 二力臂分裂成多个夹爪， 可使紧固组件中的紧固端更贴合于座舱透明件的圆弧面， 使紧固组 件在不破坏座舱透明件自身内应力的前提下， 更牢固与稳定地生成预应力来紧固座舱透明 件。
5. 本发明座舱透明件通过预应力结构与机身骨架实施紧固， 紧固组件选用具有相当强 度， 同时兼具一定弹性与韧性的材料， 当飞行过程中座舱透明件与机身骨架由于热膨胀系数 不同而产生不同程度的形变时， 可以通过紧固组件自身蕴藏的预应力补偿作用进行緩冲， 有 效化解了两者之间变形差异产生的破坏力， 保持座舱结构的安全与稳定。
6. 本发明中紧固组件选用具有相当强度， 同时兼具一定弹性与韧性的材料， 在飞行过 程中， 座舱透明件由于受到压差、 温差、 鸟撞等荷载而产生的变形与应力变化， 均可通过紧 固组件自身蕴藏的预应力的释放与再生成的过程进行緩冲，不仅不会使各种复杂的应力相互 叠加， 相反能在一定程度起到消除或减小应力集中的作用， 保持座舱结构的安全与稳定。
7. 本发明对座舱透明件与机身骨架实施紧固的过程， 不再需要在座舱透明件上开设螺 栓孔， 因此座舱透明件材质的选择将不再受紧固安装方式的困扰与限制， 相信层合无机玻璃 等以无机玻璃为主的航空玻璃组成形式在战斗机座舱透明件中的运用将会越来越广泛。
8. 本发明因为具有上述特点， 所以一体成型等复杂的座舱透明件的安装困难问题也可 得到相应地解决， 同时对座舱透明件稳定性的提高、 寿命的延长、 维修保养强度的降低等都 做出了积极的贡献。
9. 本发明预应力紧固的实施过程是通过拧紧相关螺栓来压迫紧固组件而使其产生预应 力， 在具体操作时， 通过前期的设计模块中对各个组件原材料的选择及几何形状的设计， 后 期工人只需将相关螺栓拧紧到位即可得到预设的紧固力，无须受到操作力度等不确定因素的 影响， 大大降低了操作条件和技术要求。 附图说明
图 1 为本发明战斗机座舱透明件紧固结构实施例一的釆用整块座舱透明件结构时的战 斗机整体结构示意图；
图 2为图 1的战斗机座舱正视图；
图 3为图 2的剖面图；
图 4 为本发明战斗机座舱透明件紧固结构实施例一的釆用多块座舱透明件结构时战斗 机整体结构示意图；
图 5为图 4的战斗机座舱正视图；
图 6为图 5的剖面图； 图 7、 图 23、 图 29、 图 35为图 6中的 A区域多种实施例的局部放大图； 图 8 为本发明战斗机座舱透明件紧固结构实施例一的釆用多块座舱透明件结构时战斗 机座舱侧视图；
图 9为图 8的剖面图；
图 10、 图 24、 图 30、 图 36为图 9中的 B区域多种实施例的局部放大图；
图 11、 图 25、 图 31、 图 37为图 9中的 C区域多种实施例的局部放大图；
图 12、 图 26、 图 32、 图 38为图 9中的 D区域多种实施例的局部放大图；
图 13 为本发明战斗机座舱透明件紧固结构的座舱透明件与边框的整体连接结构立体 图；
图 14、 图 21、 图 27、 图 33为本发明战斗机座舱透明件紧固结构的座舱透明件与边框 的局部连接结构多种实施例的立体图；
图 15为图 14的分解图；
图 16为本发明紧固组件的平面示意图；
图 17为本发明紧固组件的立体示意图；
图 18为本发明第二压力条的立体图；
图 19为本发明紧固组件的弧形变形区受压变形示意图；
图 20为本发明战斗机座舱透明件紧固结构中边框与座舱透明件紧固过程原理示意图； 图 22为图 21的分解图；
图 28为图 27的分解图；
图 34为图 33的分解图；
图 39为本发明战斗机座舱透明件紧固结构实施例二的战斗机座舱侧视图；
图 40为图 39的剖面图；
图 41为图 40中的 C区域的局部放大图。 具体实施方式
下面结合具体实施例对本发明作进一步说明。
请参见图 1-13 , 在本发明的第一较佳实施例中， 本发明的一种战斗机座舱透明件紧固结 构， 包括机体 1以及安装于机体 1上的座舱透明件 2; 座舱透明件 2周边结合有一边框 3。 边框 3上形成有配合机体 1的安装结构， 边框 3包括压迫组件 31和紧固组件 32, 通过压迫 组件 31和座舱透明件 2的配合压迫紧固组件 32生成预应力进而紧固座舱透明件 2。 其中战 斗机可釆用如图 1所示的釆用整块座舱透明件 2的结构；也可釆用如图 8所示的釆用多块座 舱透明件 2的结构。 请参阅图 14-17 , 为便于描述现在该实施例中作以下定义： 以图 14、 15 中座舱透明件 2的水平安装方向作为 X轴方向， 以座舱透明件 2的厚度方向作为 Y轴方向, 以座舱透明件 2的垂直安装方向作为 Z轴方向， 且 X轴垂直于所述 Y轴， Z轴垂直于 X轴 与 Y轴构成的平面； 其中：
紧固组件 32包括两个对称夹持于座舱透明件 2的弓形臂 321 , 其材料应选用具有相当 强度， 同时兼具一定弹性与韧性的材料， 如金属、 工程塑料、 高分子材料等； 两弓形臂 321 之间夹设形成一围合空间 320, 弓形臂 321包括第一力臂 3211与连接第一力臂 3211的第二 力臂 3212, 第一力臂 3211与第二力臂 3212的连接处形成一滑移端 3213 , 该滑移端 3213呈 圆弧面或斜面可以在保证在滑移过程中产生的阻力更小； 第一力臂 3211 于远离第二力臂 3212的一侧形成受压端 3214, 受压端 3214延伸形成有旋转定位棱 3217; 第二力臂 3212于 远离第一力臂 3211的一侧形成紧固端 3215, 紧固端 3215上结合有压板 3216, 且压板 3216 与第二力臂 3212的连接区域向内凹陷形成一压板位置调节区 3218, 通过该压板位置调节区 3218可在紧固过程中实现压板 3216微小的自身位置调节， 以使其更平整地贴附座舱透明件 2,第一力臂 3211的受压端 3214接受压迫组件 31的压迫并配合座舱透明件 2驱使第一力臂 3211与第二力臂 3212生成预应力。在本实施例中第一力臂 3211为一短直臂，第二力臂 3212 为一弧形臂，且第二力臂 3212的厚度自滑移端 3213至紧固端 3215形成一由厚至薄的渐变， 该种结构可以保证整个弧形臂充分和均匀形变， 不易折断。 两弓形臂 321在两受压端 3214 之间通过设置一弧形变形区 3219进行连接， 当第一力臂 3211的受压端 3214受压时， 弧形 变形区 3219 自弧形被压迫可拉伸延展， 弧形变形区 3219的受压变形过程请参阅图 19; 弧 形变形区 3219的设计保证了紧固组件 32具有一定的延展空间；紧固组件 32的两受压端 3214 之间配合形成有复数个螺栓孔。 两紧固端 3215与座舱透明件 2之间可涂抹粘结胶（如 UV 胶）或夹设双面胶（如 3M胶）或塾设緩冲垫（如橡胶片） 。 第二力臂 3212间隔形成复数 个溢流槽 32121 , 溢流槽 32121的釆用起到了紧固组件 32夹合时多余未凝固密封胶溢出作 用和防止了密封胶在凝固过程中的膨胀或收缩对紧固组件 32产生的预应力的影响， 确保紧 固组件 32与座舱透明件 2之间的预应力紧固和同时实现密封。
请参阅图 14、 15、 17, 压迫组件 31包括一第一压力条 311和一第二压力条 312; 第二 压力条 312的表面中部配合旋转定位棱 3217设置了两条通长的旋转定位槽 3121 , 该旋转定 位槽 3121的半径等于或略大于旋转定位棱 3217的半径，这样当整个透明件紧固结构分别处 于预紧固与紧固状态时，旋转定位棱 3217可以有效地在旋转定位槽 3121内定位与进行转动， 两滑移端 3213才会在第一压力条表面仅沿座舱透明件 2厚度方向位移。
第一压力条 311设置于弓形臂 321的第一力臂 3211的外侧； 紧固组件 32的弓形臂 321 的两滑移端 3213抵靠于第一压力条 311;弓形臂 321的两受压端 3214抵靠于第二压力条 312, 弓形臂 321的两紧固端 3215抵靠于座舱透明件 2连接部的两侧面。
第一压力条 311和第二压力条 312分别开设有复数个对应的螺栓孔；通过螺栓紧固第一 压力条 311与第二压力条 312; 第二压力条 312压迫弓形臂 321的两受压端 3214向第一压 力条 311方向位移， 弓形臂 321的两滑移端 3213发生相互远离的位移， 弓形臂 321的两紧 固端 3215受到座舱透明件 2的限位， 从而驱使第一力臂 3211与第二力臂 3212生成预应力 紧固座舱透明件 2。
其中第一压力条 311形成槽形孔 3111 , 并通过穿设于槽形孔 3111内的螺栓与第二压力 条 312紧固， 且第二压力条 312和紧固组件 32通过槽形孔 3111沿 Y轴方向进行位置调整； 座舱透明件 2通过围合空间 320进行 X轴方向与 Z轴方向的位置调整。 且第一压力条 311 形成压迫端 3112, 压迫端 3112的两侧形成连接端 3113 , 本实施例中连接端 3113为配合机 体 1的安装结构， 第一压力条 311通过连接端 3113 固定于机体 1 , 安装结构也可釆用其他 连接结构。
另外， 在第二力臂 3212、 座舱透明件 2以及第二压力条 312之间的空隙内填充密封胶， 从而实现更为稳定的紧固； 由于在第二力臂 3212上开设溢流槽 32121 , 溢流槽 32121的釆 用起到了紧固组件 32夹合时多余未凝固密封胶溢出作用和防止了密封胶在凝固过程中的膨 胀或收缩对紧固组件 32产生的预应力的影响。
请参阅图 7、 14-15, 当装配座舱透明件 2时， 将第二压力条 312置于紧固组件 32的围 合空间 320内并将旋转定位槽 3121与旋转定位棱 3217配合，然后在围合空间 320内填充密 封胶并将紧固组件 32的第一力臂 3211设置于第一压力条 311上， 弓形臂 321的两紧固端 3215抵靠于座舱透明件 2连接部的两侧面； 再将第一压力条 311设置于第一力臂 3211的外 侧； 弓形臂 321的两滑移端 3213抵靠于第一压力条 311 , 弓形臂 321的两受压端 3214抵靠 于第二压力条 312的外侧表面，通过依次贯穿于第一压力条、紧固组件 32和第二压力条 311 的螺栓孔的螺栓进行预紧，第二压力条 312和紧固组件 32通过槽形孔 3111沿 Y轴方向进行 位置调整； 座舱透明件 2通过围合空间 320进行 X轴方向与 Z轴方向的位置调整， 待座舱 透明件 2的位置调整到位后，通过该螺栓紧固第一压力条 311和第二压力条 312至完成紧固。
下面配合图 19来进一步说明整个紧固过程的工作原理， 弓形臂 321的两受压端 3214在 第二压力条 312的压迫作用下向第一压力条 311方向位移， 通过旋转定位棱 3217与旋转定 位槽 3121的配合保证了受压端 3214在移动过程中仅沿 Z轴方向位移，两个弓形臂 321受压 端 3214之间的距离在紧固过程中是可控（不变）的， 同时两滑移端 3213抵靠于第一压力条 311 的内侧表面发生相互远离的位移， 而两紧固端 3215沿发生相互靠近的位移直至抵靠于 座舱透明件 2的侧面， 因此两紧固端 3215的压板 3216间的距离也是可控的， 其在座舱透明 件 2上的紧固位置点也是可控的； 进一步压迫两受压端 3214向第一压力条 311方向位移， 进而驱使两滑移端 3213继续相互远离， 而两紧固端 3215此时抵靠于座舱透明件 2的侧面并 由此受到限位， 第一力臂 3211及第二力臂 3212由此发生形变并生成预应力， 至此具有稳定 预应力结构的座舱透明件 2与边框 3达到紧固状态， 座舱透明件 2获得紧固。 同样的， 当预 应力需要解除时， 只要将相应螺栓松开， 弓形臂 321的形变会恢复到之前未紧固状态， 此时 预应力自动消失， 整个战斗机座舱透明件紧固结构的部件都是无损耗的和可重复使用的， 不 仅节约了成本， 同时也非常环保。
在本实施例中，座舱透明件 2为夹层结构， 包括一内层 21和结合于内层两侧的外层 22; 外层 22形成凹陷部且外层 22边缘与内层 21平齐， 紧固端 3215贴合于凹陷部上， 此时凹陷 部为座舱透明件 2的连接部， 即外层凹陷形成连接部。
外层 22也可不形成凹陷部， 紧固端 3215直接与外层 22外表面紧密贴合， 此时连接部 为座舱透明件 2边缘， 座舱透明件 2与边框 3的连接结构请参阅图 21-26。
或者， 内层 21向外延伸形成自外层 22之间凸出的连接部； 紧固端 3215紧密贴合于内 层 21凸出外层 22的两侧面； 座舱透明件 2与边框 3的连接结构请参阅图 27-32。
另外， 座舱透明件 2的表面也可拼贴向外凸出的连接部， 在本实施例中外层 22拼贴向 外凸出的该连接部， 紧固端 3215紧密贴合于连接部的两侧面， 此时， 座舱透明件 2与边框 3的连接结构请参阅图 33-38。 请参阅图 39~41 , 在本发明的第二较佳实施例中， 本发明的一种战斗机座舱透明件紧固 结构为单块座舱透明件 2的结构， 其主要结构与第一实施例相同， 区别在于： 座舱透明件 2 内侧还拼贴有复数个沿座舱透明件 2曲面线排布于同一平面内的加强肋条 23; 加强肋条 23 的第一端贴合座舱透明件 2内表面； 加强肋条 23的第二端形成凸条 231并通过边框 3相互 紧固。
以上结合附图实施例对本发明进行了详细说明，本领域普通技术人员可根据上述说明对 本发明做出种种变化例。 因而， 实施例中的某些细节不应构成对本发明的限定， 本发明将以 所附权利要求书界定的范围作为本发明的保护范围。

Claims

权 利 要 求 书
1. 一种战斗机座舱透明件紧固结构， 包括机体以及安装于所述机体上的座舱透明件； 其特征在于， 所述座舱透明件周边结合有一边框， 所述边框上形成有配合所述机体的安装结 构， 所述边框包括压迫组件和紧固组件， 通过所述压迫组件和座舱透明件的配合压迫所述紧 固组件生成预应力进而紧固所述座舱透明件。
2. 如权利要求 1 所述的战斗机座舱透明件紧固结构， 其特征在于， 所述紧固组件包括 两个对称夹持于所述座舱透明件的弓形臂， 两弓形臂之间夹设形成一围合空间， 所述弓形臂 包括第一力臂与连接所述第一力臂的第二力臂，所述第一力臂与所述第二力臂的连接处形成 滑移端， 所述第一力臂于远离所述第二力臂的一侧形成受压端， 所述第二力臂于远离所述第 一力臂的一侧形成紧固端，所述第一力臂的受压端接受所述压迫组件的压迫并配合所述座舱 透明件驱使所述第一力臂与第二力臂生成预应力。
3. 如权利要求 2 所述的战斗机座舱透明件紧固结构， 其特征在于， 所述压迫组件包括 一第一压力条和一第二压力条；
所述第一压力条设置于所述弓形臂的第一力臂的外侧；所述紧固组件的弓形臂的两滑移 端抵靠于所述第一压力条； 所述弓形臂的两受压端抵靠于所述第二压力条， 所述弓形臂的两 紧固端抵靠于所述座舱透明件连接部的两侧面；
所述第一压力条和所述第二压力条分别开设有复数个对应的螺栓孔；通过螺栓紧固所述 第一压力条与所述第二压力条；所述第二压力条压迫所述弓形臂的两受压端向所述第一压力 条方向位移， 所述弓形臂的两滑移端发生相互远离的位移， 所述弓形臂的两紧固端受到所述 座舱透明件的限位， 从而驱使所述第一力臂与所述第二力臂生成预应力紧固所述座舱透明 件。
4. 如权利要求 3 所述的战斗机座舱透明件紧固结构， 其特征在于， 所述第一压力条形 成槽形孔， 并通过穿设于所述槽形孔内的螺栓与所述第二压力条紧固， 且所述第二压力条和 所述紧固组件通过所述槽形孔沿一第一方向进行位置调整；
所述座舱透明件通过所述围合空间进行一第二方向与一第三方向的位置调整。
5. 如权利要求 4 中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述座舱 透明件包括一内层和结合于所述内层两侧的外层； 所述外层凹陷形成所述连接部。
6. 如权利要求 4 中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述座舱 透明件包括一内层和结合于所述内层两侧的外层；所述内层向外延伸形成自所述外层之间凸 出的所述连接部。
7. 如权利要求 4 中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述座舱 透明件的表面形成向外凸出的所述连接部。
8. 如权利要求 2~7 中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述第 一力臂为一短直臂， 所述第二力臂为一弧形臂。
9. 如权利要求 2~7 中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述紧 固端上结合有压板， 且所述压板与所述第二力臂的连接区域向内凹陷形成一压板位置调节 区。
10. 如权利要求 2~7中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述两 紧固端与所述座舱透明件之间涂抹粘结胶、 夹设双面胶或塾设緩冲垫。
11. 如权利要求 2~7中任一项所述的战斗机座舱透明件紧固结构， 其特征在于， 所述第 二力臂间隔形成复数个溢流槽； 所述围合空间内填充有密封胶。
12. 如权利要求 2~7中任一项所述的战斗机座舱透明件紧固结构， 其特征在于： 所述弓 形臂的滑移端呈圆弧面或斜面。
13. 如权利要求 2~7中任一项所述的战斗机座舱透明件紧固结构， 其特征在于所述第二 力臂的厚度自所述滑移端至所述紧固端形成一由厚至薄的渐变。
14. 如权利要求 2~7中任一项所述的战斗机座舱透明件紧固结构， 其特征在于： 所述紧 固组件的弓形臂的受压端延伸形成有一旋转定位棱，所述第二压力条对应所述紧固组件的所 述旋转定位棱形成有旋转定位槽。
15. 如权利要求 2~7中任一项所述的战斗机座舱透明件紧固结构， 其特征在于： 所述弓 形臂的受压端之间通过一弧形变形区连接。
16.—种应用如权利要求 1~7中任一项紧固结构对战斗机座舱透明件进行紧固的方法。
PCT/CN2014/077241 2013-05-16 2014-05-12 战斗机座舱透明件紧固结构及其紧固方法 WO2014183612A1 (zh)

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