CN106809375A - 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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- CN106809375A CN106809375A CN201611180135.6A CN201611180135A CN106809375A CN 106809375 A CN106809375 A CN 106809375A CN 201611180135 A CN201611180135 A CN 201611180135A CN 106809375 A CN106809375 A CN 106809375A
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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, wherein the upper surface of cavity is located at rudderpost and installs on gap and separated by thermal insulation layer between rudder face, lower surface is located between internal body, and lower surface and steering wheel by phase-change material or heat sink separates；The side wall of cavity sets the conduit parallel with rudderpost bus, and capillary wick is welded on filling working medium in conduit surface, cavity, by Working fluid phase changing in cavity and diffusion, the heat of Aerodynamic Heating at rudderpost gap is dredged to the inwall of whole cavity；Rudderpost is located at internal body part and rudder face partial outer face sets thermal insulation layer.
Description
Technical field
The present invention relates to a kind of rudderpost leading-type thermal protection struc ture, belong to hypersonic aircraft thermal protection technology field.
Background technology
Hypersonic aircraft in flight course, when rudderpost be subject to direct Aerodynamic Heating when, due to rudderpost radius of curvature
It is smaller, and there is gap disturbing effect in rudder face and the gap location of body, the very little at gap of the rudderpost surface outside exposure
Region in local heat flux regions high occur, the part of rudderpost high temperature " focus " is occurred, maximum temperature possibly even surpasses
The allowable temperature of high-temperature alloy material is crossed, while local larger thermograde can cause very big thermal stress at rudderpost, it is right
The elevated temperature strength of rudderpost proposes very big challenge, and this problem is particularly acute for all movable rudder.Existing hypersonic vehicle master
The composites such as C/C, C/SiC are used to prepare rudderpost to ensure its thermal protective performance, however, there is problem of oxidation in composite,
And it is high to preparation technology requirement, cost is also higher；Metal material rudderpost has good load characteristic and the damage tolerance higher,
Reliability is high, has better performance in terms of reusable performance, but is compared with composite, and temperature in use is relatively low；It is logical
Cross the anti-heat insulation structural of composite needs to solve the thermally matched problem of different materials with the combining structure of metal connecting structure, and right
The layout and structure type of rudder have certain limitations, and cause rudder structure to complicate, and influence reliability.
The content of the invention
Technology solve problem of the invention is：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.
Technical solution of the invention is：A kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture, rudderpost is set
Hollow structure is counted into, the wherein upper surface of cavity is located at rudderpost and installs on gap and separated by thermal insulation layer between rudder face,
Lower surface is located between internal body, and lower surface and steering wheel by phase-change material or heat sink separates；The side wall of cavity is set
The conduit parallel with rudderpost bus, capillary wick is welded on conduit surface, filling working medium in cavity, by Working fluid phase changing in cavity with
Diffusion, the heat of Aerodynamic Heating at rudderpost gap is dredged to the inwall of whole cavity；Rudderpost be located at internal body part and
Rudder face partial outer face sets thermal insulation layer.
Spacing combination capillary wick between channel width w and depth h and adjacent conduit meets hair of the liquid in conduit
Resistance of the thin power more than working medium backflow.
Spacing combination capillary wick between channel width w and depth h and adjacent conduit is met in diabatic process, in rudderpost
Working medium at gap is continuous.
On the premise of structural strength and rigidity is met, channel width w is less than 0.5mm, and depth h is more than 0.4mm, and spacing is small
In 2mm.
The combining form of described capillary wick at least the two-layer capillary wick including different meshes, the wherein 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 during without conduit, according to the working environment of hypersonic aircraft, calculating meets intensity and rigidity will
The cavity wall thickness asked；
Second step, between initializing between conduit and capillary wick parameter, i.e. channel width w and depth h and adjacent conduit
Away from t, the mesh number and combining form of capillary wick；
3rd 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, according to the Temperature Distribution for obtaining, judges that hollow rudderpost is located at internal body part and positioned at rudder face portion
Whether the temperature divided meets design requirement, if meeting, current conduit and capillary wick parameter and cavity wall thickness are final setting
Meter parameter；Otherwise turn next step；
5th step, judges whether working medium of the simulation process at rudderpost gap is continuous, if continuously, reducing conduit spacing t,
Re-executed from the 3rd step；If discontinuous, reduce the mesh number of channel width w or increase capillary wick or increase capillary wick
The combination number of plies, re-executes from the 3rd step.
When conduit spacing is reduced, according to current Parameter of channel, according to the working environment of hypersonic aircraft, count again
Calculation meets the cavity wall thickness of intensity and rigidity requirement；3rd step is performed according to this cavity wall thickness.
Initiation parameter in described second step is chosen in following scope：Channel width w is less than 0.5mm, and depth h is big
In 0.4mm, 2mm, capillary wick mesh number 100-500 mesh are smaller than.
Described 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 using the hyperthermia heat pipe structure with high heat conduction ability by the local heat of heat flow province high 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),
Such that it is able to realize the isothermal of rudderpost, localized hyperthermia and the thermograde at rudderpost gap are effectively reduced so that the seam of rudderpost
Maximum temperature at gap is controlled within the reliable operation condition of metal material, while improving the thermal stress distribution of rudderpost, it is ensured that
The use intensity of rudderpost in flight course.
(2) can realize the solar heat protection of rudderpost structure, the integration of bearing function, meet low cost, high reliability, repeat and make
It is required that, with vast potential for future development.
(3) rudderpost cavity inside is using conduit and the combining form of capillary silk screen, there is provided the parallel pathways of working medium backflow, increases
The ability of strong working medium continuous backflow.
(4) it is superimposed composition sandwich construction using the different capillary wick silk screen of mesh number, increase silk screen interlayer capillary force is made
With raising working medium reflux capability.
(5) multilayer capillary core filaments web frame outer surface is the relatively low silk screen of mesh number, hardness is bigger, beneficial to wick structure and rudder
Axle cavity inner wall face is brought into close contact.
(6) alkali metal working medium is under hypersonic aircraft environmental condition, and saturated vapour pressure is relatively low, reduces rudderpost cavity
The stress level of inner structural wall face working substance steam.
(7) show that the nickel base superalloy, alkali metal and stainless steel cloth have extraordinary through research and engineering experience
Compatibility, it is ensured that heat high dredges performance.
(8) nickel base superalloy metal structure is used, good manufacturability, reliability is high.
(9) leading-type rudderpost uses heat sink or phase-change material to absorb heat positioned at internal body end face, can organize pneumatic
Heat and spread to internal body.
Brief description of the drawings
Fig. 1 is leading-type rudderpost structural representation；
Fig. 2 is that working medium flows to schematic diagram in rudderpost；
Fig. 3 a, 3b are respectively C/SiC rudderposts and leading-type rudderpost three dimension temperature distribution (unit：K)；
Fig. 4 a, 4b are respectively C/SiC rudderposts and leading-type rudderpost outer surface bus temperature distribution curve (unit：K).
Specific embodiment
Below in conjunction with the accompanying drawings and example in detail the present invention.
The leading-type rudderpost that Fig. 1 is provided for the present invention, for the connection between rudder face and body.What the present invention was provided dredges
Formula rudderpost is the hollow rudderpost of cavity type that nickel base superalloy makes, rudder face part axial wall 1., between rudder face and steering wheel at gap
Rudder wall 2. and internal body part axial wall 3. constitute high-temperature heat pipe closing housing 1, using high-temperature behavior preferably and with alkali metal work
The compatible nickel-base high-temperature alloy material of matter makes；Inner surface processes micro-channel 2 and lays capillary wick 3 and realizes capillary effect, is used for
Worker quality liquid flows back；Alkali metallic sodium 4 is filled in housing and dredges working medium as heat；Rudderpost outside wall surface and close steering wheel end face arrangement
Thermal insulation layer 5；Rudderpost is absorbed heat near 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 local Heating Characteristics high, the features such as engineering application prospect is good.
The operation principle of leading-type rudderpost is as shown in Figure 2.Under the conditions of Aerodynamic Heating, rudderpost slotted section is exposed to air
In bear local pneumatic heating, heat is conducted to inside heat pipe by tube wall, and solid-state (or liquid) alkali metal heat absorption nearby is steamed
Hair, and being flowed to unmanaged section under steam pressure effect, steam gradually lower the temperature and liquefy and be condensed into liquid by heat release, in groove and
Bringing-up section is back in the presence 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 instantiation：
Alkali-Metal Na is chosen as working medium, nickel base superalloy GH3044 is used as axial wall material, capillary wick material selection 301
Stainless steel material.
The first step, the Pneumatic pressure load in rudderpost design condition, on the premise of structural strength and rigidity is met
Choose rudderpost wall thickness parameter and Parameter of channel.It is 0.2mm to choose channel width w herein, and depth h is 0.4mm, and spacing is 2mm.
Second step, capillary wick chooses four layers of combining form, and ecto-entad is in order 100 mesh, 300 mesh, 100 mesh, 300
Mesh, wherein 300 mesh silk screens press close to conduit surface.
3rd 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, according to the Temperature Distribution for obtaining, judges that hollow rudderpost is located at internal body part and positioned at rudder face portion
Whether the temperature divided meets design requirement, if meeting, current conduit and capillary wick parameter and cavity wall thickness are final setting
Meter parameter；Otherwise turn 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 the 3rd step；If discontinuous, channel width w is reduced into 0.05mm, increase the mesh of 100 mesh capillary wicks
Number is 200 mesh, is re-executed from the 3rd 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；3rd step is performed according to this cavity wall thickness.Until the 4th step obtains final design parameter.
In order to further illustrate superiority of the invention, we have carried out conventional rudder axle construction and have existed with leading-type rudderpost structure
Temperature-responsive analysis under the conditions of identical Aerodynamic Heating.Calculation of Heat Transfer is carried out to rudderpost structure under typical thermal environment, C/SiC is obtained
Rudderpost and using dredge thermal protection struc ture rudderpost distribution of three-dimensional temperature as shown in Fig. 3 a, 3b, outer bus temperature distribution curve
As shown in Fig. 4 a, 4b.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 general knowledge as well known to those skilled in the art.
Claims (10)
1. a kind of hypersonic aircraft leading-type rudderpost thermal protection struc ture, it is characterised in that：Rudderpost is designed to hollow structure, its
The upper surface of middle cavity is located at rudderpost and installs on gap and separated by thermal insulation layer between rudder face, and lower surface is located in body
By phase-change material or heat sink separate between portion, and lower surface and steering wheel；The side wall of cavity sets parallel with rudderpost bus
Conduit, capillary wick is welded on filling working medium in conduit surface, cavity, by Working fluid phase changing in cavity and diffusion, by rudderpost gap
The heat for locating Aerodynamic Heating is dredged to the inwall of whole cavity；Rudderpost is located at internal body part and rudder face partial outer face sets
Put thermal insulation layer.
2. structure according to claim 1, it is characterised in that：Between between channel width w and depth h and adjacent conduit
Meet resistance of capillary force of the liquid in conduit more than working medium backflow away from reference to capillary wick.
3. structure according to claim 1, it is characterised in that：Between between channel width w and depth h and adjacent conduit
In diabatic process is met with reference to capillary wick, the working medium at rudderpost gap is continuous.
4. the structure according to Claims 2 or 3, it is characterised in that：On the premise of structural strength and rigidity is met, conduit
Width w is less than 0.5mm, and depth h is more than 0.4mm, is smaller than 2mm.
5. structure according to claim 1, it is characterised in that：The combining form of described capillary wick at least includes different mesh
Several two-layer capillary wicks, the wherein low capillary wick of mesh number are located at outer surface.
6. structure according to claim 1 and 2, it is characterised in that：The relative dimensions of structure are determined by following step：
The first step, it is assumed that during without conduit, according to the working environment of hypersonic aircraft, calculating meets intensity and rigidity requirement
Cavity wall thickness；
Second step, initializes conduit and capillary wick parameter, i.e. channel width w and the spacing t between depth h and adjacent conduit,
The mesh number and combining form of capillary wick；
3rd 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 axle gap, calculates the Temperature Distribution of hollow rudderpost cavity surface；
4th step, according to the Temperature Distribution for obtaining, judges that hollow rudderpost is located at internal body part and positioned at rudder face part
Whether temperature meets design requirement, if meeting, current conduit and capillary wick parameter and cavity wall thickness are final design ginseng
Number；Otherwise turn next step；
5th step, judges whether working medium of the simulation process at rudderpost gap is continuous, if continuously, reducing conduit spacing t, from the
Three steps are re-executed；If discontinuous, reduce the combination of the mesh number or increase capillary wick of channel width w or increase capillary wick
The number of plies, re-executes from the 3rd step.
7. structure according to claim 6, it is characterised in that：When conduit spacing is reduced, 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 performs the 3rd step.
8. structure according to claim 6, it is characterised in that：Initiation parameter in described second step is in following scope
Interior selection：Channel width w is less than 0.5mm, and depth h is more than 0.4mm, is smaller than 2mm, capillary wick mesh number 100-500 mesh.
9. structure according to claim 1, it is characterised in that：Described working medium uses alkali metal.
10. 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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Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107792392A (en) * | 2017-09-26 | 2018-03-13 | 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 |
| CN111428398A (en) * | 2020-03-02 | 2020-07-17 | 北京空天技术研究所 | C/SiC control surface thermal strength calculation method |
| CN111924089A (en) * | 2020-06-28 | 2020-11-13 | 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 |
| CN112357054A (en) * | 2020-11-19 | 2021-02-12 | 中国航天空气动力技术研究院 | Self-starting type heat-proof structure and high-speed aircraft |
| CN112389629A (en) * | 2020-11-19 | 2021-02-23 | 中国航天空气动力技术研究院 | Modularized wing leading edge structure and high-speed aircraft |
| CN112413100A (en) * | 2020-12-01 | 2021-02-26 | 上海航天控制技术研究所 | Long-endurance high-speed aircraft rudder shaft liquid flow cooling method and structure |
| CN112853250A (en) * | 2020-12-28 | 2021-05-28 | 哈尔滨工业大学 | Preparation method of combined gas rudder component |
| CN113665850A (en) * | 2021-08-02 | 2021-11-19 | 湖北航天技术研究院总体设计所 | Phase-change type heat-proof structure of rudder shaft and aircraft |
| CN114030589A (en) * | 2021-10-19 | 2022-02-11 | 湖北航天技术研究院总体设计所 | Light high-efficiency thermal-resistance air rudder |
| CN114180026A (en) * | 2021-12-28 | 2022-03-15 | 中南大学 | Dredging phase change composite flexible thermal protection structure and application thereof in deformable aircraft |
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Cited By (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| CN107792392A (en) * | 2017-09-26 | 2018-03-13 | 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 |
| CN111428398A (en) * | 2020-03-02 | 2020-07-17 | 北京空天技术研究所 | C/SiC control surface thermal strength calculation 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 |
| CN111924089A (en) * | 2020-06-28 | 2020-11-13 | 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 |
| CN112357054A (en) * | 2020-11-19 | 2021-02-12 | 中国航天空气动力技术研究院 | Self-starting type heat-proof 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 |
| CN112389629A (en) * | 2020-11-19 | 2021-02-23 | 中国航天空气动力技术研究院 | Modularized wing leading edge structure and high-speed aircraft |
| CN112413100A (en) * | 2020-12-01 | 2021-02-26 | 上海航天控制技术研究所 | Long-endurance high-speed aircraft rudder shaft liquid flow cooling method and structure |
| 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 |
| CN112853250A (en) * | 2020-12-28 | 2021-05-28 | 哈尔滨工业大学 | Preparation method of combined gas rudder component |
| CN113665850A (en) * | 2021-08-02 | 2021-11-19 | 湖北航天技术研究院总体设计所 | Phase-change type heat-proof structure of rudder shaft and aircraft |
| 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 |
| CN114030589A (en) * | 2021-10-19 | 2022-02-11 | 湖北航天技术研究院总体设计所 | Light high-efficiency thermal-resistance air rudder |
| CN114180026A (en) * | 2021-12-28 | 2022-03-15 | 中南大学 | Dredging phase change composite flexible thermal protection structure and application thereof in deformable aircraft |
| 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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