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Title:
AIRCRAFT PANEL, IN PARTICULAR A CABIN WINDOW, AND METHOD FOR MANUFACTURING SUCH AN AIRCRAFT PANEL
Document Type and Number:
WIPO Patent Application WO/2017/196171
Kind Code:
A1
Abstract:
The invention relates to an aircraft panel (1), in particular a cabin window. The invention also relates to an aircraft comprising at least one aircraft panel according to the invention, wherein the aircraft panel is preferably formed by a cabin window. The invention further relates to a method for manufacturing an aircraft panel, in particular a cabin window, according to the invention.

Inventors:
WIERSEMA JACOB (NL)
VAN DER SLUIS WALTER (NL)
Application Number:
PCT/NL2017/050291
Publication Date:
November 16, 2017
Filing Date:
May 10, 2017
Export Citation:
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Assignee:
AVIATION GLASS & TECH HOLDING B V (NL)
International Classes:
B64C1/14; B32B17/10
Domestic Patent References:
WO2013181484A12013-12-05
WO2014084725A12014-06-05
WO2013184904A12013-12-12
Foreign References:
US20150064374A12015-03-05
DE102014213017A12016-01-07
US20120094084A12012-04-19
Other References:
None
Attorney, Agent or Firm:
LANGENHUIJSEN, Bastiaan, Wilhelmus, Herman (NL)
Download PDF:
Claims:
Claims

1 . Aircraft panel, in particular a cabin window, comprising:

at least one ultra-thin chemically hardened first glass sheet with a maximum thickness of 1 .25 mm ;

at least one impact-absorbing intermediate layer with a maximum thickness of 0.5 mm, which intermediate layer comprises at least one polymer,

at least one polymer-comprising first fastening layer positioned between the first glass sheet and the intermediate layer for mutually fastening the first glass sheet and the intermediate layer,

at least one, preferably chemically hardened second glass sheet, which second glass sheet is positioned on a side of the intermediate layer remote from the first glass sheet, and

at least one polymer-comprising second fastening layer positioned between the second glass sheet and the intermediate layer for mutually fastening the second glass sheet and the intermediate layer,

wherein at least a part of the aircraft panel has a curved geometry.

2. Aircraft panel as claimed in claim 1 , wherein at least a part of the aircraft panel has a curved geometry obtained during the production process of the aircraft panel and/or obtained during thermoforming of the aircraft panel.

3. Aircraft panel as claimed in claim 1 or 2, wherein in non-loaded state at least a part of the aircraft panel has a curved geometry.

4. Aircraft panel as claimed in any of the foregoing claims, wherein the aircraft panel has a curved geometry.

5. Aircraft panel as claimed in any of the foregoing claims, wherein at least a part of the aircraft panel has a single curved geometry.

6. Aircraft panel as claimed in any of the foregoing claims, wherein at least a part of the aircraft panel has a multiple curved geometry.

7. Aircraft panel as claimed in any of the foregoing claims, wherein the curved geometry of at least a part of the aircraft panel has a radius lying between 1000 and 3000 millimetres. 8. Aircraft panel as claimed in any of the foregoing claims, wherein the thickness of the first glass sheet is greater than or equal to the thickness of the second glass sheet.

9. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or the second fastening layer comprise a polymer selected from the group consisting of: ethylene vinyl acetate (EVA), thermoplastic polyurethane (TPU) and polyvinyl butyral (PVB).

10. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or the second fastening layer have a maximum Young's modulus of 2100 MPa.

1 1 . Aircraft panel as claimed in any of the foregoing claims, wherein the intermediate layer comprises a polymer selected from the group of: polyethylene terephthalate (PET), polyvinyl chloride (PVC), polycarbonate (PC) and

polytetrafluoroethylene (PTFE).

12. Aircraft panel as claimed in any of the foregoing claims, wherein a surface layer of the first glass sheet and/or the second glass sheet is chemically hardened, and wherein a core layer of the first glass sheet and/or the second glass sheet takes a substantially unhardened form.

13. Aircraft panel as claimed in any of the foregoing claims, wherein the first glass sheet and/or the second glass sheet are manufactured from aluminium silicate glass or soda-lime glass.

14. Aircraft panel as claimed in any of the foregoing claims, wherein the first glass sheet has a maximum thickness of 1 .0 mm, preferably 0.7 mm, more preferably 0.55 mm, in particular 0.4 mm.

15. Aircraft panel as claimed in any of the foregoing claims, wherein the peripheral edge of the first glass sheet and/or the second glass sheet is polished.

16. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or second fastening layer have a maximum thickness of 0.4 mm, preferably 0.2 mm, more preferably 0.1 mm, in particular 0.05 mm.

17. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or the second fastening layer are transparent.

18. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or the second fastening layer have a maximum Haze of 1 % as measured in accordance with ASTM E430 and ISO 13803. 19. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or the second fastening layer are manufactured at least partially from a thermoplastic.

20. Aircraft panel as claimed in any of the foregoing claims, wherein the first fastening layer and/or the second fastening layer have an elongation at break of at least 500%, preferably at least 1 ,000%.

21 . Aircraft panel as claimed in any of the foregoing claims, wherein the thickness of the intermediate layer lies between 0.1 and 0.5 mm.

22. Aircraft panel as claimed in any of the foregoing claims, wherein the intermediate layer takes a transparent form.

23. Aircraft panel as claimed in any of the foregoing claims, wherein the intermediate layer has an elongation at break of at least 200%.

24. Aircraft panel as claimed in any of the foregoing claims, wherein the intermediate layer comprises at least one fire-retardant additive, wherein the at least one fire-retardant additive is preferably formed by an organohalogen compound.

25. Aircraft panel as claimed in any of the foregoing claims, wherein the intermediate layer has a density which is less than or equal to 1 .5 g/cm3. 26. Aircraft panel as claimed in any of the foregoing claims, wherein the intermediate layer has a maximum Young's modulus of 4150 MPa.

27. Aircraft panel as claimed in any of the foregoing claims, wherein the second glass sheet is thicker than the first glass sheet.

28. Aircraft panel as claimed in any of the foregoing claims, wherein at least one additional layer is positioned between the first glass sheet and the second glass sheet, chosen from the group consisting of: an electrochromic layer, a

thermochromic layer, a mirror layer, an opaque layer and a further glass sheet.

29. Aircraft comprising at least one at least partially curved aircraft panel as claimed in any of the foregoing claims.

30. Aircraft as claimed in claim 29, wherein the aircraft panel is received in a substantially non-loaded state in a frame.

31 . Aircraft as claimed in claim 29 or 30, wherein the aircraft panel is received in a frame such that there is a clearance between the aircraft panel and the frame. 32. Aircraft as claimed in any of the claims 29-31 , wherein at least one aircraft panel is applied as cabin window such that the first glass sheet faces toward and the second glass sheet faces away from a space enclosed by the aircraft.

33. Method for manufacturing an aircraft panel, in particular an aircraft panel as claimed in any of the claims 1 -28, comprising the steps of:

A) providing a first glass sheet with a thickness of 1 .25 mm ;

B) processing a peripheral edge of the first glass sheet;

C) chemically hardening the first glass sheet;

D) cleaning the first glass sheet;

E) positioning on the first glass sheet:

a. a polymer-comprising first fastening layer, b. an impact-absorbing intermediate layer with a maximum thickness of 0.5 mm, which intermediate layer comprises at least one polymer,

c. a polymer-comprising second fastening layer, and

d. a second glass sheet;

F) subjecting the assembly of layers formed during step E) for a time period f to an underpressure and to a temperature profile with an increased

temperature, so forming a curved laminate, and

G) allowing the laminate to cool, so forming a permanently curved aircraft

panel.

34. Method as claimed in claim 33, wherein during step F) the assembly of layers is positioned on a preferably convex bending mould. 35. Method as claimed in claim 33 or 34, wherein during step F) the assembly of layers is positioned on a convex bending mould.

36. Method as claimed in any of the claims 33-35, wherein during step F) the assembly of layers is positioned on a grating arranged on a convex bending mould.

37. Method as claimed in any of the claims 33-36, wherein during step F) the assembly of layers is covered by an air-impermeable mat provided with a valve for subjecting the assembly of layers to an underpressure. 38. Method as claimed in any of the claims 33-37, wherein during step F) the assembly of layers is subjected to a maximum temperature of between 120 and 140 degrees Celsius.

Description:
Aircraft panel, in particular a cabin window, and method for manufacturing such an aircraft panel

The invention relates to an aircraft panel, in particular a cabin window. The invention also relates to an aircraft comprising at least one aircraft panel according to the invention, wherein the aircraft panel is preferably formed by a cabin window. The invention further relates to a method for manufacturing an aircraft panel, in particular a cabin window, according to the invention. A viewing window arranged in an aircraft, also often referred to as aircraft window, is generally constructed from an outer window and a plastic inner window positioned at a distance from the outer window. The inner window is generally manufactured here from polycarbonate. A drawback of plastic inner windows is that they are relatively likely to be damaged and are often relatively dull, which obstructs light transmission and therefore has an adverse effect on the visibility through the viewing window.

A first object of the invention is to provide an improved inner window for an aircraft. A second object of the invention is to provide a relatively transparent inner window for an aircraft.

A third object of the invention is to provide a relatively scratch-resistant aircraft panel.

A fourth object of the invention is to provide a relatively scratch-resistant and/or transparent aircraft panel with a relatively good impact resistance.

A fifth object of the invention is to provide an improved inner window for an aircraft which complies with requirements set by the international aviation authorities in respect of fire resistance, self-extinguishing and smoke-generating properties.

At least one of the stated objects, or a combination thereof, can be achieved by providing an aircraft panel of the type stated in the preamble comprising: at least one ultra-thin chemically hardened first glass sheet with a maximum thickness of 1 .25 mm; at least one impact-absorbing intermediate layer with a maximum thickness of 0.5 mm, which intermediate layer comprises at least one polymer, at least one polymer-comprising first fastening layer positioned between the first glass sheet and the intermediate layer for mutually fastening the first glass sheet and the intermediate layer, at least one second glass sheet, which second glass sheet is positioned on a side of the intermediate layer remote from the first glass sheet, and at least one polymer-comprising second fastening layer positioned between the second glass sheet and the intermediate layer for mutually fastening the second glass sheet and the intermediate layer. The aircraft panel according to the invention is particularly suitable for application as inner window of an aircraft window in an aircraft. The aircraft panel according to the invention is formed by a laminate of material layers, at least one outer material layer of which is formed by a glass sheet. Glass has a significantly greater scratch-resistance than polycarbonate, this enhancing the durability of the aircraft panel. Glass is moreover a considerably clearer and more transparent material than polycarbonate, which can also enhance the transparency of the aircraft panel. The first glass sheet, and preferably also the second glass sheet, have a maximum thickness of 1 .25 mm and thereby take an ultra-thin form in order to limit the overall weight of the laminate, this being relevant for the application of the aircraft panel according to the invention in an aircraft. The weight of the aircraft panel according to the invention is preferably equal to or less than the weight of the conventional inner window manufactured from

polycarbonate. Because the first glass sheet, and preferably also the second glass sheet, take a chemically hardened and ultra-thin (<1 .25 mm) form, the first glass sheet, and preferably also the second glass sheet, will be made considerably stronger and more flexible, which considerably increases the impact resistance and considerably reduces the chance of breakage. The impact resistance of the aircraft panel as such is further increased in that the laminate moreover comprises an impact-absorbing intermediate layer which ensures that, in the case of an impact on a (chemically hardened) glass sheet, the impact is partially absorbed by

(temporary deformation of) the intermediate layer, which reduces the chance of breakage still further, this being particularly advantageous from a safety viewpoint. In the case of splintering (decomposition) of a glass sheet during an impact on the glass sheet the glass splinters will be held substantially wholly in place due to the use of the fastening layer behind, this also being particularly advantageous from a safety viewpoint. Since fire safety is also an important factor in aircraft panels, the flammable polymer-comprising intermediate layer also takes an exceptionally thin (< 0.5 mm) form in order to limit the amount of flammable material in the aircraft panel according to the invention as far as possible. The thickness of both the (flammable) first fastening layer and the (flammable) second fastening layer will generally be less than or equal to the thickness of the intermediate layer, whereby the flammability of the aircraft panel according to the invention is limited and complies with international standards and EASA regulations. In contrast to the conventional panels manufactured from polycarbonate, the above stated

embodiment of the aircraft panel according to the invention particularly complies with the Airworthiness Standards (part 25) as laid down by the Federal Aviation Administration (FAA) and with the Certification Specification -25 (CS-25) as laid down by the European Aviation Safety Agency (EASA), and more particularly with the requirements 28.853 (a), 25.853 (d), 25.775 (a) and 25.785 (d) forming part of said regulations. Regulation 28.853 (a) relates here to a plurality of fire tests, including the Ή2.5" test, the "V12" test and the "V60" test, which have been successfully completed by the above stated embodiment of the aircraft panel.

The aircraft panel according to the invention can at least partially have a

substantially planar geometry, although in the case the aircraft panel is applied as inner window of an aircraft window it will generally at least partially have a single (or multiple) curved geometry depending on the shape of the aircraft window, in particular the shape of a frame of the aircraft window configured to hold the inner window. Providing at least a part of the aircraft panel with a curved geometry is also understood to mean a bent geometry. It is preferred that at least a part of the aircraft panel has a curved geometry obtained during the production process of the aircraft panel and/or obtained during thermoforming of the aircraft panel. During production of the aircraft panel, in particular during lamination of the layers of the aircraft panel at increased temperature, the aircraft panel is therefore shaped such that the aircraft panel will be at least partially provided with a curved geometry. This curved geometry is permanent, at least in non-loaded state, unless of course the aircraft panel is once again deformed by being reheated. The curved aircraft panel also retains elasticity (flexibility) after the thermoforming, but wherein the aircraft panel will tend during deformation to return to the original, curved geometry imparted to the aircraft panel during the thermoforming. When at least a part of the aircraft panel is provided with a single curve, the relevant part of the aircraft panel, or the whole aircraft panel, is then curved relative to a transverse bending axis. It is otherwise possible here to envisage the degree of curvature varying in the direction in which the aircraft panel is curved. It is for instance possible here to envisage a lower part of the aircraft panel taking a substantially flat or lightly curved form, while an upper part of the same aircraft panel has a more pronounced curve. It is also possible to envisage the aircraft panel, or at least a part thereof, being provided with a multiple curved geometry, wherein the aircraft panel is for instance curved relative to multiple bending axes extending in the same and/or in different directions. This results in a more complex shaping of the aircraft panel. The curvature (or bending) in multiple directions can be identical, but can also differ in different directions. The curvature (radius of curvature) and/or the dimensions of the aircraft panel can vary in the longitudinal direction of the aircraft. It is therefore possible to envisage that a first aircraft panel installed in an aircraft and a second aircraft panel installed behind the first aircraft panel have different curvatures and/or dimensions. The curved geometry of at least a part of the aircraft panel according to the invention preferably has a radius lying between 1000 and 3000 millimetres, in particular between 1200 and 3000 millimetres. The (permanent) curving of at least a part of the aircraft panel in predefined manner during the production process has the advantage that the shape of the aircraft panel can be adapted to the at least partially curved framework (casing or frame) which forms part of an aircraft and in which the aircraft panel has to be installed. This has the advantage that the installation (fitting) of the aircraft panel in the framework is facilitated without the aircraft panel having to be deformed, in particular bent, during installation. The aircraft panel curved in predefined manner can therefore be installed in relatively tension-free manner. This significantly reduces the occurrence of additional stresses in the aircraft panel during installation, and thereby the risk of breakage of the aircraft panel, this considerably enhancing the safety of the aircraft panel. The aircraft panel according to the invention will generally be applied as inner window (cabin window) of an aircraft window in an aircraft. It is however also possible to envisage the aircraft panel being applied as (different type of) partition panel and/or, in the case the above stated laminate also comprises a mirror layer, as mirror in an aircraft. It is also possible to envisage applying the scratch-resistant, preferably particularly transparent, relatively fireproof, lightweight panel with relatively high impact resistance according to the invention in vessels and (land) vehicles, since aspects such as weight, scratch-resistance, transparency, fire safety and impact resistance are generally also important factors in these alternative vehicles.

As already stated, the second glass sheet is preferably also chemically hardened in order to increase the strength of the glass sheet. During the chemical hardening of the first glass sheet, and preferably also of the second glass sheet, the

(unhardened) glass is preferably immersed in a bath with molten potassium nitrate at a temperature of about 400 q C. This results in chemical exchange of K+ ions from the bath with the Na+ ions from the glass. The K+ ions (size 2.66A) take the place of the Na+ ions (size 1 .96 A) from the glass. Since they have larger dimensions they induce compressive stresses at the surface of the glass, which can thus provide more resistance. The immersion duration (also) determines the finally obtained stress level. The stress distribution does not take the same form as in the case of thermally hardened glass, and generally results in significantly stronger glass with a higher bending strength than if unhardened glass were to be hardened in thermal manner. The chemical hardening of the glass sheet can optionally take place in multiple steps, preferably in order to successively exchange different selective ions, such as sodium ions, silver ions, copper ions and/or lithium ions, with potassium ions. It is noted in this context that chemically hardened glass generally has a much higher compressive stress at the surface of the glass sheet which decreases relatively quickly just beneath the surface, wherein there is a limited tensile stress in the centre (half depth) of the glass sheet, resulting in a block-shaped stress profile. Thermally hardened glass generally has a considerably lower compressive stress at the surface of the glass sheet, wherein a relatively high tensile stress is present in the centre of the glass sheet, resulting in a parabolic stress profile. The glass applied in the glass sheet preferably comprises aluminium oxide (AI2O3), preferably in a quantity of at least 7 mol%. Such glass is also referred to as aluminium silicate glass. It has been found that, in the case of glass comprising aluminium oxide, particularly when the quantity of aluminium oxide comprises at least 7 mol%, the potassium ions (K+ ions) will penetrate deeper into the glass sheet, on average to about 50 micrometres, which imparts to the thin glass sheet a greater and thereby improved bending strength, generally of about 800 MPa. The aluminium oxide content in the glass sheet as applied in the aircraft panel according to the invention preferably lies between 7 and 25 mol%. The increased bending strength results in a relatively strong and flexible glass which has a relatively high impact resistance and which is not susceptible to vibration at all. This makes the glass sheet particularly suitable for use in and/or on an aircraft (or other type of vehicle). During curing the potassium ions will penetrate the glass sheet on two sides (on opposite (front) sides), whereby during curing potassium ions are incorporated into the glass over an overall thickness of 100 micrometres (2 x 50 micrometres). At a glass thickness of for instance 1 .0 millimetre the overall penetration depth thus amounts to 10%. A surface layer of the first glass sheet and/or the second glass sheet will thus generally be chemically hardened, and a core layer of the first glass sheet and/or second glass sheet will remain

substantially unhardened. A further advantage of applying AI2O3 in the glass sheet is that the melting temperature of the glass sheet can hereby be considerably increased, which is an additional advantage from the viewpoint of fire safety. It is also possible to envisage applying soda-lime glass instead of aluminium silicate glass. Soda-lime glass can be chemically hardened in the same manner as described above. Before chemical hardening of the first glass sheet and/or second glass sheet the peripheral edge of the first glass sheet and/or the second glass sheet is preferably polished or finished in other manner. Sharp protruding edge parts can for instance be removed by this edge finishing before the glass sheet is hardened. It is practically no longer possible, or at least not simple, to realize this edge finishing after the chemical hardening.

The first glass sheet of the aircraft panel generally faces toward people

(passengers and/or crew) present in the aircraft. The thickness of the first glass sheet is preferably greater than or equal to the thickness of the second glass sheet. The glass sheets are preferably both given an ultra-thin form (< 1 .25 mm) in order to give the glass sheets a relatively flexible character, this reducing the chance of breakage, and wherein in the case of breakage the glass sheets will break (up) into relatively small glass particles (glass splinters), this being particularly advantageous from a safety viewpoint. Such a small thickness moreover also has the advantage that the overall weight of the aircraft panel will remain limited, this being relevant and advantageous from an economic and logistical viewpoint. Furthermore, the optical deformation of the aircraft panel will in this way remain limited, which is pleasing for people looking outside via the aircraft panel. The thickness of at least one glass sheet is preferably less than 1 .25 mm, more preferably 1 .0 mm or less, in particular (about) 0.55 mm. The lower limit of the thickness of the first glass sheet and/or the second glass sheet will generally amount to 0.4 mm, in order to still be able to manufacture the glass sheet in controlled manner.

The first fastening layer and/or second fastening layer preferably has a maximum thickness of 0.4 mm. The thickness preferably amounts to a maximum of 0.2 mm, more preferably a maximum of 0.1 mm, and amounts particularly to about 0.05 mm. It is generally advantageous to apply the thinnest possible fastening layer, since this reduces the flammability of the aircraft panel. The first fastening layer, and preferably also the second fastening layer preferably have a maximum Young's modulus of 2100 MPa. Because of this low Young's modulus the fastening layers have a relatively flexible character, this being advantageous in absorbing an impact exerted on the aircraft panel. The first fastening layer and/or the second fastening layer preferably have an elongation at break of at least 500%, preferably at least 1 ,000%. Such a high elongation at break prevents the fastening layers cracking relatively quickly, whereby the aircraft panel as such can be kept intact for longer.

In a preferred embodiment the first fastening layer and/or the second fastening layer comprise a thermoplastic polymer. The first fastening layer and/or the second fastening layer preferably comprise a polymer selected from the group consisting of: ethylene vinyl acetate (EVA), thermoplastic polyurethane (TPU) and polyvinyl butyral (PVB). EVA is generally preferred for the intended application in an aircraft, since EVA is the least flammable of the above stated polymers. In the case TPU is applied as polymer in the first fastening layer and/or second fastening layer, it is then preferred, if the panel is applied as (inner) window, to apply an aliphatic TPU since this is substantially colourless, and, if the panel is not applied as window but as other type of partition panel and/or mirror, it is then generally preferred to apply an aromatic TPU. An aromatic TPU has a yellow colour but is less flammable than an aliphatic TPU. It is possible here to envisage the intermediate layer

(intentionally) being provided with a colour, whereby the panel can for instance also be applied as reflective panel.

The first fastening layer and/or the second fastening layer are preferably substantially wholly transparent. It is advantageous in the case of the fastening layers when the Haze is less than or equal to 1 % as measured in accordance with ASTM E430 and ISO 13803. The Haze is indicative of the measure of transparency and relates particularly to an (undesirable) optical effect caused by microscopic structures or a residue on a surface (measured in accordance with ASTM E430 and ISO 13803).

The impact-absorbing intermediate layer preferably comprises a thermoplastic polymer. The impact-absorbing intermediate layer preferably comprises a polymer selected from the group consisting of: polyethylene terephthalate (PET), polyvinyl chloride (PVC), polycarbonate (PC) and polytetrafluoroethylene (PTFE). Such thermoplastics are substantially transparent and have the property of being able to absorb an impact relatively well, whereby these polymers are particularly suitable for application as or in an intermediate layer. In order to bring about improved fastening of the intermediate layer to the adjacent layers, generally formed by the first fastening layer and the second fastening layer, it is generally advantageous to subject the intermediate layer to a surface treatment before laminating the intermediate layer with the other material layers to form the actual aircraft panel. This surface treatment can for instance be formed by a plasma treatment. In order to reduce the flammability of the intermediate layer (and/or at least one fastening layer), this being advantageous from a safety viewpoint, it is preferred that the intermediate layer (and/or at least one fastening layer) comprises at least one fire- retardant additive. This additive prevents the spread of fire or a least counters spread of fire. The additive is preferably formed by an organohalogen compound. Such compounds are able to remove reactive H- and OH-radicals during a fire. The organohalogen compound preferably comprises bromine and/or chlorine. Preferred from a viewpoint of fire retardance over an organochlorine compound such as PCB (polychlorinated biphenyl) is an organobromine compound such as PBDE

(polybrominated diphenyl ether). Other examples of applicable brominated compounds are: Tetrabromobisphenol A, Decabromodiphenyl ether (Deca),

Octabromodiphenyl ether, Tetrabromodiphenyl ether, Hexabromocyclododecane (HBCD), Tribromophenol, Bis(tribromophenoxy)ethane, Tetrabromobisphenol A polycarbonate oligomer (TBBA or TBBPA), Tetrabromobisphenol A epoxy oligomer (TBBA or TBBPA), and Tetrabromophthalic acid anhydride. Other examples of applicable chlorinated compounds are: chlorinated paraffin, Bis(hexachlorocyclopentadieno)cyclooctane, Dodecachloride pentacyclodecane (Dechlorane), and 1 ,2,3,4,7,8,9, 10, 13, 13, 14, 14-dodecachloro- 1 ,4,43,5,6,68,7, 10, 10a, 1 1 , 12, 12a-dodecahydro-1 ,4,7, 10- dimethanodibenzo[a,e]cyclooctene (Dechlorane Plus). From an environmentally- friendly and economic viewpoint however, it is generally preferred not to apply an organohalogen compound in the intermediate layer. Such an intermediate layer without organohalogen compound which is sufficiently fireproof can be realized by giving the intermediate layer a sufficiently thin form, preferably by adhering to a maximum thickness of 150 micrometres.

The thickness of the intermediate layer preferably lies between 0.1 and 0.5 mm. A thicker intermediate layer is undesirable from a fire safety viewpoint. When the intermediate layer is given a fire-retardant form, for instance by adding one or more of said additives, the intermediate layer can then be given a thicker form than if the intermediate layer were not given a fire-retardant form. The thickness of a fire- retardant intermediate layer preferably lies between 200 and 500 micrometres. The thickness of an intermediate layer which is not fire-retardant preferably lies between 100 and 150 micrometres. The intermediate layer preferably has an elongation at break of at least 200%, this being sufficient to be able to absorb a considerable impact. The intermediate layer preferably has a density which is lower than or equal to 1 .5 g/cm 3 . Materials with a higher density make the aircraft panel undesirably heavy. The intermediate layer preferably has a maximum Young's modulus of 4150 MPa. The intermediate layer will generally be (slightly) more rigid here than the first fastening layer and second fastening layer, but will still be just flexible enough to be sufficiently able to absorb an impact. It is also possible to envisage applying a glass-comprising intermediate layer instead of a polymer-comprising intermediate layer. The glass-comprising intermediate layer can optionally be formed here by an optionally chemically hardened glass sheet. The maximum thickness of this intermediate (third) glass sheet is preferably also 1 .25 mm, more preferably 1 .0 mm, and in particular about 0.55 mm.

It is conceivable, and can be advantageous, for at least one additional layer to be positioned between the first glass sheet and the second glass sheet, preferably chosen from the group consisting of: an electrochromic layer, a thermochromic layer, a mirror layer, an opaque layer and a further glass sheet. The mirror layer can take diverse forms. It is possible here to envisage the mirror layer being embodied as a film reflective on at least one side. An advantage of a film is that the layer thickness of the mirror layer is substantially homogenous, which can enhance homogenous reflection of the mirror. It is also possible to envisage a (thin) metal (oxide) layer being arranged on another layer of the laminate, this other carrier layer preferably being formed by the glass sheet. Examples of suitable metals are copper, silver, gold, nickel, aluminium, Beryllium, chrome, molybdenum, platinum, rhodium, tungsten and titanium. The metal layer can be arranged on the carrier layer, in particular the glass sheet, by means of vacuum vapour deposition techniques and/or sputtering. The arranged metal layer can optionally be at least partially removed, for instance by means of sandblasting, in order to make a part of the mirror wholly or semi-transparent and/or to impart a satinized (matt)

appearance to the mirror. This makes it possible to generate visual effects behind the mirror layer, for instance in a separate material layer, which will be visible via the semi-transparent mirror to persons looking in the mirror. The above stated examples of the mirror layer are embodiments wherein the (static) mirror layer takes a permanently specular form. In a preferred embodiment a side of the mirror layer remote from the glass sheet is at least partially provided with a coating which protects the mirror layer. The coating is particularly advantageous when the mirror layer is formed by a metal layer so that oxidation of the metal layer can be prevented or at least countered. If the mirror layer is formed by a copper layer, it is for instance possible to envisage covering the copper layer with an inhibitor on the basis of for instance azole derivative. It is however also possible to envisage the mirror layer taking a semi-permanent (temporarily) specular form. The mirror layer can generally be made specular as desired here. This is possible for instance by having at least a part of the mirror layer formed by an electrochromic layer.

Connecting the electrochromic layer, optionally on the basis of liquid crystals (LCD), to an electrical energy source such as a battery enables the layer to be charged, whereby the specular layer can be activated or deactivated. The electrochromic layer can optionally be co-laminated during the production process. Later assembly of such a layer with the already formed laminate can also be envisaged. It is possible to envisage positioning the thermochromic layer behind an optionally non-specular, optionally made non-specular, part of the mirror, particularly of the glass sheet. The aircraft panel according to the invention can be provided with a ventilation opening which passes through the aircraft panel. The ventilation opening connects here to an outer side of the first glass sheet and an outer side of the second glass sheet.

The invention also relates to a vehicle, in particular an aircraft, comprising at least one aircraft panel according to the invention. At least one aircraft panel is preferably applied here as cabin window, more preferably such that the first glass sheet faces toward, and the second glass sheet faces away from a space enclosed by the aircraft. The aircraft panels can optionally also be applied as partition panel, video screen, touchscreen, mirror and/or combinations thereof. Vehicles are understood to mean, among others, motorbikes, automobiles, vessels and airborne vehicles (aircraft). The at least one aircraft panel preferably takes an at least partially curved form, as already described at length in the foregoing. The aircraft panel is preferably received in a substantially non-loaded state in a frame or framework. This means that the aircraft panel is not, or at least not appreciably deformed by the framework and is substantially only held in position by the framework. The aircraft panel is preferably received in a frame such that there is clearance between the aircraft panel and the frame. The clearance ensures that the aircraft panel is movable to a limited extent - generally in the order of magnitude of 1 -2 millimetres - relative to the frame. This generally enhances the substantially tension-free retention of the aircraft panel.

The invention also relates to a method for manufacturing an aircraft panel according to the invention, comprising the steps of: A) providing a first glass sheet with a thickness of 1 .25 mm; B) processing a peripheral edge of the first glass sheet; C) chemically hardening the first glass sheet; D) cleaning the first glass sheet; E) positioning on the first glass sheet: a polymer-comprising first fastening layer, an impact-absorbing intermediate layer with a maximum thickness of 0.5 mm, which intermediate layer comprises at least one polymer, a polymer-comprising second fastening layer and a second glass sheet; F) subjecting the assembly of layers formed during step E) for a time period f to an underpressure and to a temperature profile with an increased temperature, so forming a laminate. The time period t preferably lies between 90 and 270 minutes. The method preferably also comprises step G), comprising of allowing the formed laminate to cool, so forming a permanently curved aircraft panel. The temperature is preferably increased stepwise during step F), more preferably by at least 40 degrees per step until a final temperature is reached. A temperature step can for instance be applied every 30 minutes. A typical maximum temperature lies above the glass transition

temperature of the polymer materials applied, and will generally lie between 100 and 200 degrees Celsius, and preferably lies between 120 and 140 degrees Celsius. A temperature profile is preferably applied during step F) with an average rise in temperature of about 1 degree Celsius per minute. It can be advantageous during step F), when the maximum temperature has been reached, to gradually relieve the underpressure. Further advantages and embodiment variants of the obtained laminated aircraft panel according to the invention have already been described at length in the foregoing. The bending (curving) of at least a part of the laminate generally takes place in step F). It is otherwise also possible to envisage having the bending (curving) of at least a part of the laminate take place after manufacture of the laminate, so after step F), as post-processing step. Generally however it is strongly preferred from an efficiency viewpoint to have the curving process take place during the laminating step as according to step F). It is advantageous here during step F) to position the assembly of layers on a

preferably convex bending mould. The convex bending mould is provided with a convex (spherical) upper side, generally with a curvature which also has to be arranged in the aircraft panel to be formed. The convex bending mould can for instance be manufactured from aluminium or another suitable metal. It can be advantageous to position a grating, in particular a metal grating, provided with openings on top of the convex bending mould, whereby the desired underpressure can be realized in improved manner. The grating can be given a flexible form, whereby the grating takes on the same curvature as the convex upper side of the bending mould. It is also possible to envisage the grating being given a

substantially rigid form and being provided with convex shaping that has to be arranged in the aircraft panel to be formed. It is possible to envisage the grating forming part of the bending mould, wherein the grating can optionally be arranged releasably on another part of the bending mould. The air passage openings in the grating have a typical cross-section of between 3 and 7 millimetres. The assembly of layers for laminating, which at that moment will generally still have a substantially flat (uncurved) geometry, will be placed on top of the grating. The assembly of layers is preferably covered during step F) by an air-impermeable mat provided with a valve for creating an underpressure between the mat and the underlying bending mould, and the optionally applied grating, whereby the assembly of layers is subjected as such to a desired underpressure. The weight of the mat will moreover bend the assembly of layers over the convex surface of the underlying structure (grating and/or bending mould).

Non-limitative embodiment variants of the invention are described in the clauses below: 1 . Aircraft panel, in particular a cabin window, comprising:

at least one ultra-thin chemically hardened first glass sheet with a maximum thickness of 1 .25 mm ;

at least one impact-absorbing intermediate layer with a maximum thickness of 0.5 mm, which intermediate layer comprises at least one polymer, ■ at least one polymer-comprising first fastening layer positioned between the first glass sheet and the intermediate layer for mutually fastening the first glass sheet and the intermediate layer,

at least one, preferably chemically hardened second glass sheet, which second glass sheet is positioned on a side of the intermediate layer remote from the first glass sheet, and

at least one polymer-comprising second fastening layer positioned between the second glass sheet and the intermediate layer for mutually fastening the second glass sheet and the intermediate layer. 2. Aircraft panel according to clause 1 , wherein the thickness of the first glass sheet is greater than or equal to the thickness of the second glass sheet.

3. Aircraft panel according to clause 1 or 2, wherein the first fastening layer and/or the second fastening layer comprise a polymer selected from the group consisting of: ethylene vinyl acetate (EVA), thermoplastic polyurethane (TPU) and polyvinyl butyral (PVB).

4. Aircraft panel according to any of the foregoing clauses, wherein the first fastening layer and/or the second fastening layer have a maximum Young's modulus of 2100 MPa. 5. Aircraft panel according to any of the foregoing clauses, wherein the intermediate layer comprises a polymer selected from the group of: polyethylene terephthalate (PET), polyvinyl chloride (PVC), polycarbonate (PC) and

polytetrafluoroethylene (PTFE).

6. Aircraft panel according to any of the foregoing clauses, wherein a surface layer of the first glass sheet and/or the second glass sheet is chemically hardened, and wherein a core layer of the first glass sheet and/or the second glass sheet takes a substantially unhardened form.

7. Aircraft panel according to any of the foregoing clauses, wherein the first glass sheet and/or the second glass sheet are manufactured from aluminium silicate glass or soda-lime glass.

8. Aircraft panel according to any of the foregoing clauses, wherein the first glass sheet has a maximum thickness of 1 .0 mm, preferably 0.7 mm, more preferably 0.55 mm, in particular 0.4 mm. 9. Aircraft panel according to any of the foregoing clauses, wherein the peripheral edge of the first glass sheet and/or the second glass sheet is polished.

10. Aircraft panel according to any of the foregoing clauses, wherein the first fastening layer and/or second fastening layer have a maximum thickness of 0.4 mm, preferably 0.2 mm, more preferably 0.1 mm, in particular 0.05 mm.

1 1 . Aircraft panel according to any of the foregoing clauses, wherein the first fastening layer and/or the second fastening layer are transparent. 12. Aircraft panel according to any of the foregoing clauses, wherein the first fastening layer and/or the second fastening layer have a maximum Haze of 1 % as measured in accordance with ASTM E430 and ISO 13803.

13. Aircraft panel according to any of the foregoing clauses, wherein the first fastening layer and/or the second fastening layer are manufactured at least partially from a thermoplastic.

14. Aircraft panel according to any of the foregoing clauses, wherein the first fastening layer and/or the second fastening layer have an elongation at break of at least 500%, preferably at least 1 ,000%.

15. Aircraft panel according to any of the foregoing clauses, wherein the thickness of the intermediate layer lies between 0.1 and 0.5 mm. 16. Aircraft panel according to any of the foregoing clauses, wherein the intermediate layer takes a transparent form.

17. Aircraft panel according to any of the foregoing clauses, wherein the intermediate layer has an elongation at break of at least 200%.

18. Aircraft panel according to any of the foregoing clauses, wherein the intermediate layer comprises at least one fire-retardant additive, wherein the at least one fire-retardant additive is preferably formed by an organohalogen compound.

19. Aircraft panel according to any of the foregoing clauses, wherein the intermediate layer has a density which is less than or equal to 1 .5 g/cm 3 .

20. Aircraft panel according to any of the foregoing clauses, wherein the intermediate layer has a maximum Young's modulus of 4150 MPa.

21 . Aircraft panel according to any of the foregoing clauses, wherein the second glass sheet is thicker than the first glass sheet. 22. Aircraft panel according to any of the foregoing clauses, wherein at least one additional layer is positioned between the first glass sheet and the second glass sheet, chosen from the group consisting of: an electrochromic layer, a thermochromic layer, a mirror layer, an opaque layer and a further glass sheet. 23. Aircraft panel according to any of the foregoing clauses, wherein the aircraft panel takes a curved form.

24. Aircraft comprising at least one aircraft panel according to any of the foregoing clauses.

25. Aircraft according to clause 24, wherein at least one aircraft panel is applied as cabin window such that the first glass sheet faces toward and the second glass sheet faces away from a space enclosed by the aircraft.

26. Method for manufacturing an aircraft panel according to any of the clauses 1 -23, comprising the steps of:

A) providing a first glass sheet with a thickness of 1 .25 mm ;

B) processing a peripheral edge of the first glass sheet;

C) chemically hardening the first glass sheet;

D) cleaning the first glass sheet;

E) positioning on the first glass sheet:

a. a polymer-comprising first fastening layer, b. an impact-absorbing intermediate layer with a maximum

thickness of 0.5 mm, which intermediate layer comprises at least one polymer,

c. a polymer-comprising second fastening layer, and

d. a second glass sheet;

F) subjecting the assembly of layers formed during step E) for a time period f to an underpressure and to a temperature profile with an increased

temperature, so forming a laminate.

The invention will be elucidated on the basis of non-limitative exemplary

embodiments shown in the following figures. Herein:

figure 1 is a schematic side view of an aircraft panel according to the present invention;

figure 2 shows schematically an aircraft panel embodied as cabin window according to the present invention; and

figure 3 shows schematically a method for manufacturing an aircraft panel according to the present invention. Figure 1 is a schematic side view of an aircraft panel (1 ) according to the invention. Panel (1 ) comprises a first ultra-thin chemically hardened glass sheet (2) with a maximum thickness of 1 .25 mm and a second glass sheet (3) which is optionally also chemically hardened. The glass is for instance an aluminosilicate, which material is hard and not likely to break. Such a glass comprises for instance 15- 25% AI2O3 and 50-60% S1O2. The glass can also comprise 15-35% earth-alkaline metal (beryllium, magnesium, calcium, strontium, barium, radium) in order to improve chemical hardening of the glass. The glass has for instance a Young's modulus between 60 and 80 GPa.

Chemical hardening of glass sheet (2) ensures that the glass becomes

exceptionally strong, and also ensures for instance that, if the glass sheet were to break, forming of splinters is prevented and/or that the formed glass splinters are relatively small. The chemical hardening takes place for instance by immersing glass sheet (2) in a bath of molten salt at a high temperature, for instance of about 400 degrees Celsius. Possible steps for processing glass sheet (2) preferably take place before the chemical hardening, since hardening of the glass impedes further processing of the glass. Following chemical hardening glass sheet (2) is for instance washed, for instance by ultrasonic washing, so that glass sheet (2) is clean and can adhere better to adjacent material layers.

Situated between the two glass sheets (2, 3) is a polymer intermediate layer (4) with a maximum thickness of 0.5 mm. Intermediate layer (4) has an impact- absorbing effect. Intermediate layer (4) is for instance made from fire-retardant polyethylene terephthalate (PET), polyvinyl chloride (PVC), polycarbonate (PC) or polytetrafluoroethylene (PTFE or Teflon). Intermediate layer (4) is preferably transparent so that the panel can be applied as transparent window. Intermediate layer (4) is preferably also fire-retardant, for instance due to halogens being applied in intermediate layer (4). The thickness of the intermediate layer generally lies between 100 and 500 micrometres. The density of intermediate layer (4) generally lies between 0.9 and 1 .5 grams per cubic centimetre. In an alternative embodiment intermediate layer (4) is a central glass sheet (4).

Intermediate layer (4) is connected to the two glass sheets (2, 3) by means of two polymer fastening layers (5, 6). Fastening layers (5, 6) are for instance made from ethylene vinyl acetate (EVA), thermoplastic polyurethane (TPU) or polyvinyl butyral (PVB). The thickness of fastening layers (5, 6) is preferably as thin as possible, and less than 0.38 mm, in particular less than 0.2 mm. Fastening layer (5, 6) is generally relatively flammable, wherein reducing the thickness, and so the amount of fastening material, limits the fuel available to a possible fire. A fastening layer (5, 6) of EVA has for instance a Young's modulus between 0.01 and 0.05 GPa. EVA is relatively poorly flammable compared to TPU and PVB, and so is a preferred material for a fire-retardant aircraft panel (1 ). Fastening layers (5, 6) are preferably also transparent for use of the panel in a cabin window. The fastening layers have for this purpose a low opacity so that it is at least 95%, in particular at least 99% light-permeable.

Figure 2 shows schematically a part of an aircraft (1 1 ) provided with an aircraft panel (12), or cabin window (12), according to the present invention. The panel will generally be given a curved form here. Additional advantages of the panel applied according to the invention, in addition to being light in weight and having a relatively high impact resistance, are having a relatively homogenous light permeability, the high measure of scratch-resistance and having a uniform thickness, whereby the light refraction is likewise relatively uniform. Aircraft panel (12) is optionally provided with a ventilation opening (not shown) which passes through aircraft panel (12).

Figure 3 shows schematically a method for manufacturing an aircraft panel (21 ) according to the present invention as also shown in figures 1 and 2. The method has the steps of:

A) providing a first glass sheet (22) with a thickness (d) of 1 .25 mm;

B) processing a peripheral edge (27) of the first glass sheet (22);

C) chemically hardening the first glass sheet (22), for instance by immersing glass sheet (22) in a bath of molten salt;

D) cleaning the first glass sheet (22);

E) positioning on the first glass sheet (22) a polymer-comprising first fastening layer (25), an impact-absorbing intermediate layer (24) with a maximum thickness of 0.5 mm, which intermediate layer (24) comprises at least one polymer, a polymer- comprising second fastening layer (26), and a second glass sheet (23) ; F) subjecting the formed assembly of layers (22-26) for a time period f to an underpressure (P) and to a temperature profile (T) with an increased temperature, so forming a laminate. The processing of peripheral edge (27) of first glass sheet (22) takes place before first glass sheet (22) is chemically hardened. This is because glass sheet (22) is easier to process before it is hardened. Cleaning of first glass sheet (22) takes place for instance by placing the glass sheet in an ultrasonic bath, wherein sound waves are transmitted through the bath. Ultrasonic cleaning of glass sheet (22) has the advantage that no scratches occur on glass sheet (22), that the whole glass sheet (22) is cleaned and that the cleaning can take place relatively quickly.

Subjecting the assembly of layers (22-26) to an underpressure (P) takes place for instance under vacuum. The temperature (T) to which the assembly of layers (22- 26) is subjected is for instance about 105 degrees Celsius.

Putting together the assembly of layers (22-26) and/or subjecting the assembly to an underpressure and increased temperature take place for instance in a clean room. A temperature of between 18 and 23 degrees Celsius is preferably maintained in such a room, with a relative air humidity of about 20%. A reliable lamination of the material layers takes place in such conditions.

During step F) the assembly of layers (22-26) can be curved and formed until a desired shape is achieved. This has the advantage that the finally preformed aircraft panel is fully ready for use and can be enclosed practically without further deformation and relatively tension-free in a framework arranged in an aircraft, this considerably increasing the durability, strength, reliability (predictability) and ease of installation of aircraft panel (21 ) compared to conventional elastic panels which are deformed during installation in the framework until a desired curved geometry is obtained. A further detail view of the curving of the assembly of layers (22-26) during step F) is shown in figure 3G. Figure 3G more specifically shows a vacuum oven (30) which encloses a receiving space (31 ), the temperature T and pressure P of which can be regulated. Installed in this receiving space (31 ) is a bending mould (32) manufactured from aluminium and having a convex upper surface (32a).

Arranged on the convex upper surface (32a) of bending mould (32) is a flexible grating (33), also referred to as grid or wire mesh, which is provided with air passage openings (not shown), wherein each of the air passage openings has a typical cross-section of between 0.15 and 0.35 mm 2 . Grating (33) can for instance be manufactured from metal and/or plastic. Grating (33) can optionally be deemed as part of bending mould (32). The assembly of layers (22-26), which at that moment still does not have a permanent curved geometry, is placed on top of grating (33). Some curvature will generally occur however because of the own weight of the assembly of layers (22-26). The assembly of layers (22-26) has a smaller footprint than grating (33), whereby grating (33) protrudes all around relative to the assembly of layers (22-26). The assembly of layers (22-26) is then covered by a substantially air-impermeable mat (34), in particular a rubber mat, provided with an underpressure valve (35). Due to the weight of mat (34) the assembly of glass layers (22-26) is bent further, and possibly already substantially wholly over the convex surface of grating (33). The footprint of mat (34) is larger here than that of grating (33). Underpressure valve (35) is located a distance from the assembly of layers (22-26), but connects to an upper side of grating (33). By applying an underpressure, in particular a vacuum, in vacuum oven (30) air will escape via underpressure valve (35), wherein substantially all air around the assembly of layers (22-26) will also be extracted. This underpressure results in the assembly of layers (22-26) being pressed firmly over the curved upper side of grating (33), whereby the assembly takes on the desired, curved shape. Layers (22-26) will be connected to each other by sufficiently increasing the temperature in vacuum oven (30), generally in stepwise manner, to for instance about 130 degrees Celsius. The actual curved aircraft panel (21 ) will be preformed by subsequent controlled cooling.

It will be apparent that the invention is not limited to the exemplary embodiments shown and described here, but that within the scope of the appended claims numerous variants are possible which will be self-evident to the skilled person in this field. It is possible here to envisage that different inventive concepts and/or technical measures of the above described embodiment variants can be wholly or partially combined without departing from the inventive concept described in the appended claims. The verb "comprise" and conjugations thereof used in this patent specification are understood to mean not only "comprise", but also the terms "consist of",

"substantially consist of", "formed by" and conjugations thereof.