CN106809375B - A kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture - Google Patents
A kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture Download PDFInfo
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- CN106809375B CN106809375B CN201611180135.6A CN201611180135A CN106809375B CN 106809375 B CN106809375 B CN 106809375B CN 201611180135 A CN201611180135 A CN 201611180135A CN 106809375 B CN106809375 B CN 106809375B
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- rudderpost
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- capillary wick
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C9/00—Adjustable control surfaces or members, e.g. rudders
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
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Abstract
A kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture, rudderpost is designed to hollow structure, the upper surface of its hollow cavity is located on rudderpost installation gap and is separated between rudder face by thermal insulation layer, lower surface is located at internal body, and by phase-change material or heat sink separates between lower surface and steering engine；The conduit parallel with rudderpost bus is arranged in the side wall of cavity, and capillary wick is welded on conduit surface, fills working medium in cavity, by Working fluid phase changing in cavity and diffusion, the heat of Aerodynamic Heating at rudderpost gap is dredged to the inner wall of entire cavity；Rudderpost is located at internal body part and rudder face partial outer face setting thermal insulation layer.
Description
Technical field
The present invention relates to a kind of rudderpost leading-type thermal protection struc tures, belong to hypersonic aircraft thermal protection technology field.
Background technique
Hypersonic aircraft is in flight course, when rudderpost is by direct Aerodynamic Heating, due to rudderpost radius of curvature
It is smaller, and in the gap location of rudder face and body there are gap disturbing effect, rudderpost surface outside the exposure very little at gap
Region in will appear local high heat flux regions, so that the part of rudderpost high temperature " hot spot " is occurred, maximum temperature possibly even surpasses
The allowable temperature of high-temperature alloy material is crossed, while the biggish temperature gradient in part can cause very big thermal stress at rudderpost, it is right
The elevated temperature strength of rudderpost proposes very big challenge, this problem is particularly acute for all movable rudder.Existing hypersonic vehicle master
The composite materials such as C/C, C/SiC are used to prepare rudderpost to guarantee its thermal protective performance, however, composite material is there are problem of oxidation,
And height is required to preparation process, cost is also higher；There is metal material rudderpost good load characteristic and higher damage to extend the deadline,
High reliablity has better performance in terms of reusable performance, but compares with composite material, lower using temperature；It is logical
The composite structure for crossing the anti-heat insulation structural of composite material and metal connecting structure needs to solve the problems, such as the thermally matched of different materials, and right
The layout and structure type of rudder have certain limitations, and cause rudder structure to complicate, influence reliability.
Summary of the invention
Technology of the invention solves the problems, such as: to solve the deficiencies in the prior art, the present invention provides a kind of superb
Velocity of sound aircraft leading-type rudderpost thermal protection struc ture.
The technical solution of the invention is as follows: a kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture sets rudderpost
Hollow structure is counted into, the upper surface of hollow cavity is located on rudderpost installation gap and is separated between rudder face by thermal insulation layer,
Lower surface is located at internal body, and by phase-change material or heat sink separates between lower surface and steering engine；The side wall of cavity is arranged
The conduit parallel with rudderpost bus, capillary wick are welded on conduit surface, and working medium is filled in cavity, by Working fluid phase changing in cavity with
Diffusion, the heat of Aerodynamic Heating at rudderpost gap is dredged to the inner wall of entire cavity；Rudderpost be located at internal body part and
Thermal insulation layer is arranged in rudder face partial outer face.
Spacing combination capillary wick between channel width w and depth h and adjacent conduit meets hair of the liquid in conduit
Thin power is greater than the resistance of working medium reflux.
Spacing combination capillary wick between channel width w and depth h and adjacent conduit meets in diabatic process, in rudderpost
Working medium at gap is continuous.
Under the premise of meeting structural strength and rigidity, channel width w is less than 0.5mm, and depth h is greater than 0.4mm, and spacing is small
In 2mm.
The combining form of the capillary wick includes at least two layers of capillary wick of different meshes, wherein the low capillary wick of mesh number
Positioned at outer surface.
The relative dimensions of structure are determined by following step:
The first step, it is assumed that when without conduit, according to the working environment of hypersonic aircraft, calculating meets intensity and rigidity is wanted
The cavity wall thickness asked；
Second step initializes conduit and capillary wick parameter, i.e. between channel width w and depth h and adjacent conduit between
Away from t, the mesh number and combining form of capillary wick；
Third step, foundation meets conduit and the hollow rudderpost heat and mass transfer analysis model of capillary wick parameter is emulated, root
According to the aerodynamic loading condition at rudderpost gap, the Temperature Distribution of hollow rudderpost cavity surface is calculated；
4th step judges that hollow rudderpost is located at internal body part and is located at rudder face portion according to obtained Temperature Distribution
Whether the temperature divided meets design requirement, if satisfied, then current conduit and capillary wick parameter and cavity wall thickness are final set
Count parameter；Otherwise turn in next step；
5th step judges whether working medium of the simulation process at rudderpost gap is continuous, if continuously, reducing conduit spacing t,
It is re-executed from third step；If discontinuous, reduce channel width w or increase the mesh number of capillary wick or increase capillary wick
The number of plies is combined, is re-executed from third step.
When reducing conduit spacing, counted again according to current Parameter of channel according to the working environment of hypersonic aircraft
Calculate the cavity wall thickness for meeting intensity and rigidity requirement；Third step is executed according to this cavity wall thickness.
Initiation parameter in the second step is chosen in the following range: channel width w is less than 0.5mm, and depth h is big
In 0.4mm, spacing is less than 2mm, capillary wick mesh number 100-500 mesh.
The working medium uses alkali metal.
Rudderpost cavity material selects nickel base superalloy, capillary wick material selection stainless steel material.
The present invention has the beneficial effect that compared with prior art
(1) present invention is using having the hyperthermia heat pipe structure of high thermal conductivity ability for the local heat of high heat flow province at rudderpost gap
Amount is quickly dredged to other large area regions and heat sink structure of rudderpost tube wall (or being absorbed by internal body phase-change material),
So as to realize the isothermal of rudderpost, the localized hyperthermia at rudderpost gap and temperature gradient is effectively reduced, so that the seam of rudderpost
Maximum temperature at gap controls within the reliable operation condition of metal material, while improving the thermal stress distribution of rudderpost, guarantees
The use intensity of rudderpost in flight course.
(2) it can be realized the solar heat protection of rudderpost structure, the integration of bearing function, meet low cost, high reliability, repeat and make
It is required that having vast potential for future development.
(3) rudderpost cavity inside uses the combining form of conduit and capillary silk screen, provides the parallel pathways of working medium reflux, increases
The ability of strong working medium continuous backflow.
(4) it is superimposed using the different capillary wick silk screen of mesh number and forms multilayered structure, increased silk screen interlayer capillary force and make
With raising working medium reflux capability.
(5) multilayer capillary core filaments web frame outer surface is the lower silk screen of mesh number, and hardness is larger, is conducive to wick structure and rudder
Axis cavity inner wall face fits closely.
(6) for alkali metal working medium under hypersonic aircraft environmental condition, saturated vapour pressure is lower, reduces rudderpost cavity
The stress level of inner structural wall face working substance steam.
(7) through research and engineering experience, to show that the nickel base superalloy, alkali metal and stainless steel cloth have extraordinary
Compatibility, it is ensured that high heat dredges performance.
(8) nickel base superalloy metal structure, good manufacturability, reliability height are used.
(9) leading-type rudderpost is located at internal body end face and uses heat sink or phase-change material to absorb heat, can organize pneumatic
It heats and is spread to internal body.
Detailed description of the invention
Fig. 1 is leading-type rudderpost structural schematic diagram；
Fig. 2 is that working medium flows to schematic diagram in rudderpost；
Fig. 3 a, 3b are respectively C/SiC rudderpost and leading-type rudderpost three dimension temperature distribution (unit: K)；
Fig. 4 a, 4b are respectively C/SiC rudderpost and leading-type rudderpost outer surface bus temperature distribution curve (unit: K).
Specific embodiment
With reference to the accompanying drawing and example in detail is of the invention.
Fig. 1 is leading-type rudderpost provided by the invention, for the connection between rudder face and body.It is provided by the invention to dredge
Formula rudderpost is the hollow rudderpost of cavity type of nickel base superalloy production, rudder face part axial wall 1., between rudder face and steering engine at gap
Rudder wall 2. and internal body part axial wall 3. constitute high-temperature heat pipe closing shell 1, using high-temperature behavior preferably and with alkali metal work
The compatible nickel-base high-temperature alloy material production of matter；Inner surface processing micro-channel 2 is simultaneously laid with the realization capillary effect of capillary wick 3, is used for
Worker quality liquid reflux；Alkali metallic sodium 4 is filled in shell as heat dredges working medium；Rudderpost outside wall surface and close steering engine end face are arranged
Thermal insulation layer 5；Rudderpost is absorbed heat close to internal body end face using heat sink or phase-change material 6.Compared with traditional passive type thermal protection struc ture,
The heat structure has lightweight, samming, high-strength, adapts to the high Heating Characteristics in part, the features such as engineering application prospect is good.
The working principle of leading-type rudderpost is as shown in Figure 2.Under the conditions of Aerodynamic Heating, rudderpost slotted section is exposed to air
Middle receiving local pneumatic heating, heat are conducted by tube wall to inside heat pipe, and the heat absorption of neighbouring solid-state (or liquid) alkali metal is steamed
Hair, and flowing under steam pressure effect to unmanaged section, gradually heat release cools down and liquefies and is condensed into liquid steam, in groove and
Bringing-up section is back under the action of capillary wick internal capillary forces.By the vaporization heat absorption of alkali metal working medium and liquidation exothermic reaction, realize
Fire end heat reduces heating end temperature and bulk temperature gradient to the fast transport at non-heated end.
Leading-type rudderpost design cycle is discussed below by way of specific example:
Alkali-Metal Na is chosen as working medium, nickel base superalloy GH3044 is as axial wall material, capillary wick material selection 301
Stainless steel material.
The first step, according to the Pneumatic pressure load in rudderpost design condition, under the premise of meeting structural strength and rigidity
Choose rudderpost wall thickness parameter and Parameter of channel.Choosing channel width w herein is 0.2mm, and depth h is 0.4mm, spacing 2mm.
Second step, capillary wick choose four layers of combining form, and ecto-entad is 100 mesh, 300 mesh, 100 mesh, 300 in order
Mesh, wherein 300 mesh silk screens are close to conduit surface.
Third step, foundation meets conduit and the hollow rudderpost heat and mass transfer analysis model of capillary wick parameter is emulated, root
According to the aerodynamic loading condition at rudderpost gap, the Temperature Distribution of hollow rudderpost cavity surface is calculated；
4th step judges that hollow rudderpost is located at internal body part and is located at rudder face portion according to obtained Temperature Distribution
Whether the temperature divided meets design requirement, if satisfied, then current conduit and capillary wick parameter and cavity wall thickness are final set
Count parameter；Otherwise turn in next step；
5th step judges whether working medium of the simulation process at rudderpost gap is continuous, if continuously, conduit spacing t is subtracted
Small 0.2mm, re-executes from third step；If discontinuous, channel width w is reduced into 0.05mm, increases the mesh of 100 mesh capillary wicks
Number is 200 mesh, is re-executed from third step.
6th step, according to parameter current, according to the working environment of hypersonic aircraft, recalculate meet intensity and just
Spend desired cavity wall thickness；Third step is executed according to this cavity wall thickness.Until the 4th step obtains final design parameter.
Superiority in order to further illustrate the present invention, we have carried out conventional rudder axle construction and leading-type rudderpost structure exists
Temperature-responsive analysis under the conditions of identical Aerodynamic Heating.Calculation of Heat Transfer is carried out to rudderpost structure under typical thermal environment, obtains C/SiC
Rudderpost and use dredge the distribution of three-dimensional temperature of the rudderpost of thermal protection struc ture as shown in Fig. 3 a, 3b, outer bus temperature distribution curve
As shown in Fig. 4 a, 4b.As it can be seen that the maximum temperature and localized temperature gradients of rudderpost can be substantially reduced using thermal protection struc ture is dredged,
With good anti-thermal effect.
Unspecified part of the present invention belongs to common sense well known to those skilled in the art.
Claims (8)
1. a kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture, it is characterised in that: rudderpost is designed to hollow structure,
The upper surface of hollow cavity is located on rudderpost installation gap and is separated between rudder face by thermal insulation layer, and lower surface is located in body
Portion, and by phase-change material or heat sink separated between lower surface and steering engine；The side wall setting of cavity is parallel with rudderpost bus
Conduit, capillary wick are welded on conduit surface, working medium are filled in cavity, by Working fluid phase changing in cavity and diffusion, by rudderpost gap
The heat of place's Aerodynamic Heating is dredged to the inner wall of entire cavity；Rudderpost is located at internal body part and rudder face partial outer face is set
Set thermal insulation layer；Spacing combination capillary wick between channel width w and depth h and adjacent conduit meets hair of the liquid in conduit
Thin power is greater than the resistance of working medium reflux；The relative dimensions of structure are determined by following step:
The first step, it is assumed that when without conduit, according to the working environment of hypersonic aircraft, calculate and meet intensity and rigidity requirement
Cavity wall thickness；
Second step, initialization conduit and capillary wick parameter, i.e. spacing t between channel width w and depth h and adjacent conduit,
The mesh number and combining form of capillary wick；
Third step, foundation meets conduit and the hollow rudderpost heat and mass transfer analysis model of capillary wick parameter is emulated, according to rudder
Aerodynamic loading condition at axis gap calculates the Temperature Distribution of hollow rudderpost cavity surface；
4th step judges that hollow rudderpost is located at internal body part and positioned at rudder face part according to obtained Temperature Distribution
Whether temperature meets design requirement, if satisfied, then current conduit and capillary wick parameter and cavity wall thickness are that final design is joined
Number；Otherwise turn in next step；
5th step judges whether working medium of the simulation process at rudderpost gap is continuous, if continuously, reducing conduit spacing t, from
Three steps re-execute；If discontinuous, reduce channel width w or increase the mesh number of capillary wick or increase the combination of capillary wick
The number of plies is re-executed from third step.
2. structure according to claim 1, it is characterised in that: between channel width w and depth h and adjacent conduit
Away from meeting in diabatic process in conjunction with capillary wick, the working medium at rudderpost gap is continuous.
3. structure according to claim 1 or 2, it is characterised in that: under the premise of meeting structural strength and rigidity, conduit
Width w is less than 0.5mm, and depth h is greater than 0.4mm, and spacing is less than 2mm.
4. structure according to claim 1, it is characterised in that: the combining form of the capillary wick includes at least different mesh
Two layers several of capillary wicks, wherein the low capillary wick of mesh number is located at outer surface.
5. structure according to claim 1, it is characterised in that: when reducing conduit spacing, according to current Parameter of channel, root
According to the working environment of hypersonic aircraft, the cavity wall thickness for meeting intensity and rigidity requirement is recalculated；According to this cavity wall
Thickness executes third step.
6. structure according to claim 1, it is characterised in that: the initiation parameter in the second step is in following range
Interior selection: channel width w is less than 0.5mm, and depth h is greater than 0.4mm, and spacing is less than 2mm, capillary wick mesh number 100-500 mesh.
7. structure according to claim 1, it is characterised in that: the working medium uses alkali metal.
8. structure according to claim 1, it is characterised in that: rudderpost cavity material selects nickel base superalloy, capillary wick
Material selection stainless steel material.
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| CN201611180135.6A CN106809375B (en) | 2016-12-19 | 2016-12-19 | A kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture |
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| CN201611180135.6A CN106809375B (en) | 2016-12-19 | 2016-12-19 | A kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture |
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| CN106809375B true CN106809375B (en) | 2019-05-24 |
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| CN107792392B (en) * | 2017-09-26 | 2019-07-12 | 北京航天长征飞行器研究所 | A kind of active complement heat conduction pilot system of flying vehicles control rudder tipping leading edge and method |
| CN111428398B (en) * | 2020-03-02 | 2023-10-13 | 北京空天技术研究所 | C/SiC control surface thermal strength calculation method |
| CN111924089B (en) * | 2020-06-28 | 2021-09-07 | 北京临近空间飞行器系统工程研究所 | Rudder shaft heat-proof structure with separated heat-proof and force-bearing functions |
| CN112389629B (en) * | 2020-11-19 | 2022-10-18 | 中国航天空气动力技术研究院 | Modularized wing leading edge structure and high-speed aircraft |
| CN112357054B (en) * | 2020-11-19 | 2022-06-24 | 中国航天空气动力技术研究院 | Self-starting type heat-proof structure and high-speed aircraft |
| CN112413100B (en) * | 2020-12-01 | 2022-07-01 | 上海航天控制技术研究所 | Long-endurance high-speed aircraft rudder shaft liquid flow cooling method and structure |
| CN112853250B (en) * | 2020-12-28 | 2022-08-05 | 哈尔滨工业大学 | Preparation method of combined gas rudder component |
| CN113665850B (en) * | 2021-08-02 | 2023-06-13 | 湖北航天技术研究院总体设计所 | Phase-change heat-proof structure of rudder shaft and aircraft |
| CN114030589B (en) * | 2021-10-19 | 2023-07-21 | 湖北航天技术研究院总体设计所 | Light high-efficiency thermal resistance air rudder |
| CN114180026B (en) * | 2021-12-28 | 2023-12-01 | 中南大学 | Composite flexible heat protection structure for dredging phase change and application of composite flexible heat protection structure in deformable aircraft |
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