KR20230170458A - Aircraft anti-icing or de-icing system and method making the same - Google Patents

Abstract

A carbon fiber stack layer in which several plies made of the carbon fiber material of the present invention are stacked; A heating element that supplies heat to remove ice that forms on the aircraft surface; The invention relates to an anti-icing and de-icing system for an aircraft wing and a method of manufacturing the system, which is formed of a heat conductor layer that conducts heat generated from the heating element; and an insulating layer that prevents heat generated from the heating element from being conducted into the aircraft interior.

Description

Aircraft anti-icing or de-icing system and method making the same}

The present invention relates to an aircraft anti-icing and de-icing system and a method of manufacturing the same, and more specifically, to prevent ice formation on the surface of the aircraft due to changes in local airflow conditions around the aircraft, using CFRP, aluminum foil and cork board. It relates to an aircraft anti-icing and de-icing system in which a heating element is formed between laminated structures and a method of manufacturing the same.

Freezing is a very threatening factor in terms of aircraft stability. It occurs when supercooled liquid droplets collide with the surface of the aircraft while moving through the air in areas of low atmospheric temperature and high relative humidity. Ice that occurs on the surface of an aircraft wing can change the surface shape of the wing and become a factor that impairs aerodynamic characteristics such as reduced lift. Additionally, freezing that occurs around the engine intake can cause flow separation or intake air reduction, thereby reducing engine performance.

In particular, it was revealed that the cause of the crash of Air France Flight 447, which occurred on June 1, 2009, was the accumulation of ice crystals on the pitot probe, so aircraft icing plays a very important role in the safety and life of passengers.

In order to prevent aircraft performance deterioration and major accidents caused by icing, various devices to prevent icing (e.g. high-temperature compressed air from turbine engines, pneumatic boot expansion, structural vibration) are installed in areas with a high risk of icing, such as wings. impulse or deformation, etc.) is installed. The device is called an Ice Protection System (IPS), and the IPS is mainly placed on the leading edge of the aircraft's wing.

However, the high-temperature compressed air method is not suitable for CFRP structures, which have low thermal conductivity and are vulnerable to high temperatures, and the pneumatic boot expansion and structural impact methods require a minimum thickness of ice accumulation on the wings, so the anti-icing function is limited.

As carbon-fiber reinforced polymer (CFRP) is widely used in aircraft structures, it is necessary to be more careful when designing IPS due to its relatively low thermal conductivity and its vulnerability to high temperatures compared to existing aluminum.

In order to solve the above problem, the present invention is an aircraft room where a heating element is formed between structures laminated with CFRP, aluminum foil, and cork board to prevent ice formation on the surface of the aircraft due to changes in local airflow conditions around the aircraft. It relates to an ice-making system and a method of manufacturing the same.

The anti-icing and de-icing system for the aircraft wing according to an embodiment of the present invention includes a carbon fiber stack layer in which several plies made of carbon fiber materials are laminated, a heating element that supplies heat to remove ice formed on the surface of the aircraft, and a heating element. It may be formed of a heat conductor layer that conducts heat generated and an insulation layer that prevents heat generated from the heating element from being conducted into the aircraft interior.

In this embodiment, the carbon fiber stack layer has a heat conductor layer laminated on one side and an insulation layer on the other side, the insulation layer is in contact with the aircraft, and the heating element is between the heat conductor layer and the carbon fiber stack layer. It may be characterized as being laminated in at least one of several layers of plies made of a fiber material, or between a carbon fiber stack layer and an insulation layer.

In this embodiment, the heat conductor layer and the insulation layer may be respectively bonded to one side and the other side of the carbon fiber stack layer with epoxy.

In this embodiment, the carbon fiber stack layer may be characterized as being laminated in multiple layers of plies formed of carbon fiber reinforced polymer.

In this embodiment, the heating element is characterized in that it is made of heat-resistant High Performance Plastic that is formed to surround the outside of the electric heating element and insulate the electric current generated in the electric heating element from the outside. You can.

In this embodiment, the heating element may be formed along the longitudinal direction of the aircraft wing, and may be provided in plural pieces spaced apart at regular intervals in the width direction of the wing.

In this embodiment, the heat conductor layer may be formed of aluminum.

In this embodiment, the insulation layer may be formed of cork material.

A method of manufacturing an anti-icing and de-icing system for an aircraft wing according to an embodiment of the present invention includes the steps of forming a carbon fiber stack layer by stacking several plies made of carbon fiber material, and forming ice on the surface of the aircraft in the carbon fiber stack layer. It can be formed by attaching a heating element that supplies heat to remove , and inserting epoxy into the carbon fiber stack layer to which the heating element is attached and curing it.

In this embodiment, the step of laminating a heat conductor layer that conducts heat generated from the heating element on one surface of the carbon fiber stack layer may be further included.

In this embodiment, the step of laminating an insulating material layer on the other side of the carbon fiber stack layer to prevent heat generated from the heating element from being conducted into the aircraft interior may be further included.

In this embodiment, the step of laminating at least one of the heat conductor layer and the insulation layer and bonding them together under vacuum pressure using epoxy may be further included.

In this embodiment, in the step of attaching a heating element that supplies heat to the carbon fiber stack layer to remove ice formed on the surface of the aircraft, the heating element is used to insulate the electric heating element and the electric current formed in the electric heating element from the outside. It is made of heat-resistant high-performance plastic formed to surround the exterior of the electrothermal heating element, and may further include the step of plasma surface treatment of the surface of the high-performance plastic.

The present invention is more suitable for the IPS structure of existing aircraft using CFRP.

Additionally, the icing phenomenon that occurs on the surface of the aircraft can be controlled more precisely through the heating element.

Additionally, since parts already installed on the aircraft can be used, the burden of additional costs is reduced.

Additionally, it is highly durable as it can operate without high temperatures or shocks that affect the aircraft.

The effects that can be obtained from the present invention are not limited to the effects mentioned above, and other effects not mentioned can be clearly understood by those skilled in the art from the description below. will be.

Figure 1 is a cross-sectional view showing the materials and stacking sequence of an aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 2 is a cross-sectional view showing the positions of heating elements in an aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 3 is a cross-sectional view showing that the aircraft anti-icing and de-icing system according to an embodiment of the present invention is formed along the longitudinal direction of the aircraft wing and is provided with a plurality of units spaced apart at regular intervals in the width direction of the wing.
Figure 4 is a block showing a method of manufacturing an aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 5 is a cross-sectional view showing the presence or absence of materials, the stacking order, and the position of the heating element in the aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 6 is a graph showing the temperature change on the surface where the heating element is located according to the embodiment and comparative example of the aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figures 7a and 7b are graphs showing temperature distribution according to distance from the center of the heating element in the panel of the example and comparative example of the aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 8 is a diagram showing the heat distribution in a steady state for Examples 2 and 3 of the aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 9a is a diagram showing the results of simulation of airflow streamlines and ice formation formed by 15 heating elements at 2 mm intervals in an aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 9b is a diagram showing airflow streamlines formed by five heating elements 30 at 50 mm intervals and the results of icing simulation in an aircraft anti-icing and de-icing system according to an embodiment of the present invention.
Figure 10 shows simulation results formed by 15 heating elements spaced at 2 mm intervals in an aircraft anti-icing and de-icing system according to an embodiment of the present invention, (a) airfoil cross-section temperature contour, (b) surface temperature when the angle of attack is 0°, ( c) This is a diagram showing the surface temperature when the angle of attack is 4°.
Figure 11 shows simulation results formed by five heating elements 30 at 50 mm intervals in an aircraft anti-icing and de-icing system according to an embodiment of the present invention, (a) airfoil cross-section temperature contour, (b) surface when the angle of attack is 0°. Temperature, (c) This is a diagram showing the surface temperature when the angle of attack is 4°.

The terms used in this specification will be briefly explained, and the present invention will be described in detail.

The terms used in the present invention are general terms that are currently widely used as much as possible while considering the function in the present invention, but this may vary depending on the intention or precedent of a person working in the art, the emergence of new technology, etc. In addition, in certain cases, there are terms arbitrarily selected by the applicant, and in this case, the meaning will be described in detail in the description of the relevant invention. Therefore, the terms used in the present invention should be defined based on the meaning of the term and the overall content of the present invention, rather than simply the name of the term.

When it is said that a part "includes" a certain element throughout the specification, this means that, unless specifically stated to the contrary, it does not exclude other elements but may further include other elements. Additionally, throughout the specification, “A or B,” “at least one of A and B,” “at least one of A or B,” “A, B, or C,” “at least one of A, B, and C,” and “ Each phrase such as "at least one of A, B, or C" means that it may include any one of the items listed together in that phrase, or any possible combination thereof.

Below, with reference to the attached drawings, embodiments of the present invention will be described in detail so that those skilled in the art can easily implement the present invention. However, the present invention may be implemented in many different forms and is not limited to the embodiments described herein. In order to clearly explain the present invention in the drawings, parts that are not related to the description are omitted, and similar parts are given similar reference numerals throughout the specification.

Hereinafter, various embodiments of the present invention are described with reference to the attached drawings.

Referring to FIG. 1, the anti-icing and de-icing system 1 for the aircraft wing according to one embodiment includes a carbon fiber stack layer 20, a heating element 30, a heat conductor layer 10, and an insulation layer 40. Corresponds to the cross-sectional view.

The carbon fiber stack layer 20 is formed in a structure in which plies made of carbon fiber materials are stacked in multiple layers. Carbon fiber is a material with many advantages such as high rigidity, high tensile strength, high strength-to-weight ratio, high chemical resistance and high temperature resistance, and is widely used in general areas such as sports and civil engineering, as well as military and aerospace fields.

However, carbon fiber reinforced polymer is a material widely used in aircraft structures, but it has a relatively low thermal conductivity and is more vulnerable to high temperatures than aluminum, so special care needs to be taken when designing anti-icing and de-icing systems.

Additionally, as an example, a carbon fiber stack may be formed by stacking multiple plies, which are prepregs made of carbon fiber material. Here, carbon fiber can be used as carbon-fiber reinforced polymer (CFRP).

The heating element 30 is formed as an element that supplies heat to remove ice formed on the surface of the aircraft. In the conventional IPS system, when high-temperature air from a turbine engine is used, pneumatic boot expansion, and structural vibration or deformation are performed, installation is limited due to CFRP, which is vulnerable to high temperatures, and the minimum amount of ice accumulation.

The heating element 30, especially the electrothermal IPS, has the feature of being able to be controlled to precisely heat specific surfaces of the aircraft based on the Joule heating principle. In addition, compared to conventional technologies, it has the advantage of being easy to install and operating energy-efficiently.

Additionally, the heating element 30 may be composed of an electric heating element 31 and high-performance plastic 32. The electrothermal heating element 31 is formed of a film-shaped metal foil, so that structural integrity can be maintained inside the anti-icing and de-icing system 1 of the aircraft wing. In one embodiment, the metal foil may be a thin etched metal foil. High performance plastic (32) refers to a material that has excellent mechanical, chemical, and thermal stability compared to general plastic. In particular, when the high-performance plastic 32 is made of polyimide, it has excellent insulation and heat resistance, preventing electrical leakage occurring in the electric heating element 31 and maintaining its shape at high temperatures.

The heat conductor layer 10 corresponds to a layer that conducts heat generated from the heating element 30 to the surface of the aircraft. In other words, since carbon fiber reinforced polymer has low thermal conductivity, the heat generated from the heating element 30 needs to be evenly transmitted to the surface of the aircraft in order to prevent icing from occurring on the surface of the aircraft. In particular, when the heating elements 30 are located at regular intervals, it is further necessary to reduce the temperature drop in areas where the heating elements 30 are not located.

Additionally, the heat conductor layer 10 may be formed of aluminum foil, which is an aluminum material. Aluminum foil has better thermal conductivity than carbon fiber reinforced polymer, and can even play a role in preventing corrosion when attached to the surface of an aircraft.

The insulation layer 40 is formed in a structure to prevent heat generated from the heating element 30 from being conducted into the aircraft interior. In order for the heat generated from the heating element 30 to be effectively transferred to the surface of the aircraft, it is necessary to block it from being transferred to the interior of the aircraft. In addition, when the heating elements 30 are placed on the surface of the aircraft at regular intervals, it is necessary to reduce the temperature drop so that the parts where the heating elements 30 are not located do not freeze.

Additionally, the insulation layer 40 may use a cork board made of cork material. Cork has low thermal conductivity and low density, so it is suitable for insulation and has the advantage of being light in weight.

Referring to Figures 1 and 2, according to one embodiment, the carbon fiber stack layer 20 has a heat conductor layer 10 laminated on one side and an insulation layer 40 on the other side, and the insulation layer 40 is In contact with the aircraft, the heating element 30 is between the heat conductor layer 10 and the carbon fiber stack layer 20, between several layers of plies made of carbon fiber material, or between the carbon fiber stack layer 20 and the insulation layer 40. It can be stacked in at least one of the spaces. That is, the heat conductor layer 10 - the carbon fiber stack layer 20 - the insulation layer 40 are laminated in the order, and the heating element 30 is the heat conductor layer 10 as shown in (a) to (c) of FIG. 2. and the insulation layer 40. The heating element 30 may be located on one side or the other side of the carbon fiber stack layer 20, or between several layers of plies made of carbon fiber material. This involves cutting the carbon fiber stack to the size of the heating element 30 and installing it on that portion. Alternatively, it can be installed by attaching it to the carbon fiber stack layer (20).

Additionally, referring to FIG. 2, the heat conductor layer 10 and the insulation layer 40 may be respectively adhered to one side and the other side of the carbon fiber stack layer 20 with epoxy 50. Since carbon fiber stacks, heat conductors, and insulation materials are different materials, they are not easy to join due to their different thermal properties. In particular, when attached to the surface of an aircraft, thermal stability and physical properties must be increased, so each material can be joined with epoxy (50) resin.

1 to 3 show a heating element 30 located on the anti-icing and de-icing system 1 of the aircraft wing, according to an embodiment. The location, spacing, and number of heating elements 30 are particularly important for the efficiency of de-icing and de-icing of aircraft wings.

In particular, referring to Figure 2, since the thermal conductivity of aluminum foil, which is a heat conductor, is higher than that of carbon fiber reinforced polymer, the efficiency of the system may vary depending on which layer it is located. In addition, the closer the heating element 30 is to the heat conductor, the more heat is transferred to the surface of the aircraft, resulting in greater thermal benefits. However, to prevent the heating element 30 from being imprinted on the aluminum foil, an adhesive bond line between the heating element 30 and the aluminum foil is used. Be careful to keep it thin.

Also, referring to FIG. 3, the heating elements 30 are formed along the longitudinal direction (x-axis) of the aircraft wing, and a plurality of heating elements 30 may be provided at regular intervals in the width direction (y-axis) of the wing. The heating element 30 can be installed on the entire surface of the aircraft wing, but in the case of aircraft wings, air currents are formed locally depending on the shape, and the method of heat transfer is different for each location, so controlling them individually can improve energy efficiency. The heating element 30 can be installed appropriately depending on the shape and size of the wing. However, when the heating element 30 is located at regular intervals on the aircraft wing, icing may occur due to a temperature drop in the area where the heating element 30 is not located, so attention must be paid to the size and spacing of the heating element.

Referring to FIG. 4, the manufacturing method (S1) of the anti-icing and de-icing system 1 for the aircraft wing according to one embodiment includes the step of forming a carbon fiber stack layer 20 by stacking several plies made of carbon fiber material. (S10), attaching a heating element 30 that supplies heat to the carbon fiber stack layer 20 to remove ice formed on the surface of the aircraft (S20), and a carbon fiber stack layer to which the heating element 30 is attached ( It can be manufactured by inserting epoxy 50 into 20) and curing it (S30).

In particular, in the step (S20) of attaching the heating element 30 that supplies heat to the carbon fiber stack layer 20 to remove ice formed on the surface of the aircraft, the heating element 30, as described above, is attached to the carbon fiber stack layer ( It can be located on one side, the other side, or between several layers of plies made of carbon fiber material, and this can be done by cutting the carbon fiber stack to the size of the heating element 30 and installing it on that part, or on the carbon fiber stack layer 20. It can be installed by attaching it.

The step (S30) of inserting and curing epoxy 50 into the carbon fiber stack layer 20 to which the heating element 30 is attached is efficient because the carbon fiber stack and the heating element 30 are not easy to join due to their different characteristics. To join, epoxy (50) resin can be injected to join. In one embodiment, the carbon fiber stack and the heating element 30 can be sealed in a vacuum bag and cured in an autoclave oven at a maximum temperature of 120°C at 6 atm.

In addition, a step (S40) of laminating a heat conductor layer 10 that conducts heat generated from the heating element 30 on one surface of the carbon fiber stack layer 20 may be further included. Heat generated from the heating element 30 is efficiently transferred to the surface of the aircraft without exposing the heating element 30 to the outside to protect against freezing, and it can be laminated on the top of the carbon fiber stack layer 20 to prevent corrosion. In one embodiment, the heat conductor layer 10 may be washed with acetone and then co-cured with the carbon fiber stack layer 20.

In addition, a step (S50) of laminating an insulation layer 40 on the other surface of the carbon fiber stack layer 20 to prevent heat generated from the heating element 30 from being conducted into the aircraft interior may be further included. As described above, the insulation layer 40 is installed to block heat generated from the heating element 30 from being transferred to the interior of the aircraft and to prevent freezing due to temperature drop in areas where the heating element 30 is not located.

Also, in one embodiment, in order to attach the heat conductor layer 10 and the insulation layer 40 to the carbon fiber stack layer 20, a room temperature curing epoxy (50) adhesive paste may be used to attach them under vacuum pressure. After attaching the carbon fiber stack layer 20 and the heating element 30, attaching the heat conductor layer 10 and the insulation layer 40 maintains the initial thickness of the insulation material and prevents residual wrapping due to asymmetric stacking. This is to prevent.

In addition, in the step (S20) of attaching the heating element 30 that supplies heat to the carbon fiber stack layer 20 to remove ice formed on the surface of the aircraft, the heating element 30 is connected to the electric heating element 31 and the electric heat generating element 31. It is made of heat-resistant, high-performance plastic (32) formed to surround the outside of the electric heating element (31) in order to insulate the current generated in the heating element (31) from the outside, and the surface of the high-performance plastic (32) is subjected to plasma surface treatment. Step (S21) may be further included. Plasma surface treatment is a method used for antifouling, improvement of adhesion, micro-patterning to miniaturize surface patterns, anti-fingerprint surface treatment, etc. When plasma surface treatment of the surface of high-performance plastic (32) is performed, the adhesion with epoxy (50) is improved. can be further improved.

In the aircraft wing part anti-icing and de-icing system (1), the type, characteristics, dimensions, and arrangement intervals of each material can be changed in various ways depending on the aircraft's wing shape, dimensions, aircraft operating position and speed, etc. In addition, the aircraft wing part de-icing and de-icing system (1) of the present invention is used not only for fixed-wing aircraft such as general passenger aircraft or fighter jets, but also for helicopters, unmanned aerial vehicles (UAVs), urban air mobility (UAM), etc. It can also be applied to devices. The effect according to the present invention can be judged through actual manufacturing examples and comparative examples below.

[Manufacture and thermal response of the anti-icing and de-icing system of the aircraft wing]

Example

The present invention was tested for effective ice removal by setting parameters in the anti-icing and de-icing system (1) of the aircraft wing according to one embodiment. Here, the parameters correspond to the presence or absence of layers with different materials, the stacking order of each material, and the location and number of heating elements.

The carbon fiber (CFRP) used in the experiment was carbon/epoxy unidirectional lamina prepreg (USN125-B, SK chemicals), the heat conductor was aluminum foil (1100-H19 alloy), and the insulation was cork board. Additionally, epoxy adhesive paste (Loctite EA 9394 AERO, Henkel Corporation Aerospace) was used to bond the heat conductor and the insulation material.

The dimensions of the insulation system used in the experiment were 330mm The heating element 30 is an electrothermal heating film 31 (model KHA-112/10, OMEGA Engineering) formed of a thin etched metal foil between two polyimide layers, and was formed with a width of 25 mm, a length of 305 mm, and a thickness of 0.114 mm.

Referring to Figure 5, the panel forming the system of Example 1 is made by sequentially stacking 0.1 mm aluminum foil, 0.96 mm CFRP, and 4.6 mm cork board, and applying room temperature curing epoxy adhesive paste to the aluminum foil and cork board. It was glued using. The heating element 30 was located between CFRP and the cork board, and the distance between the heating elements 30 was 50 mm. Voltage was applied to the electrothermal heating element 31 through a direct current power supply.

Comparison example

The present invention presents an anti-icing and de-icing system (1) for aircraft wings according to a comparative example. Referring to Figure 5, in Comparative Examples 1 to 4, the heating element 30 is located in the second layer from the top among the CFRP layers. In Comparative Examples 2 to 4, aluminum foil with a thickness of 0.1 mm was laminated, and in Comparative Examples 3 and 4, a 4.6 mm thick cork board was laminated. (However, when the entire material was co-cured in Comparative Examples 3 and 4, due to the curing cycle It was observed that the cork board collapsed permanently to a thickness of 1.7 mm.) In Comparative Example 4, the spacing between the heating elements 30 was set to 75 mm.

Other than this, the configuration was the same as in the embodiment.

Referring to FIG. 6, it is a graph showing the temperature change at the upper surface where the heating element 30 is located over time. The example is maintained at a lower temperature than Comparative Example 1, but higher than Comparative Examples 2 to 4. This indicates that the presence of the aluminum foil and the cork board can maintain the temperature of the surface of the aircraft high through the aluminum foil using the heat generated from the heating element 30 with the same power.

[Multiphysics anti-icing simulation in the normal state of the anti-icing and de-icing system of the aircraft wing]

Examples and comparative examples

The present invention presents an experimental comparison of the thermal characteristics of the system through steady-state thermal finite element modeling (FEM) of the embodiment and comparative example in the anti-icing and de-icing system 1 of the aircraft wing according to an embodiment.

In this simulation, the dimensions of the system in the stacked structure of Example 1 and Example 1 are expanded to 600 mm, and the heating element 30 in the stacked structure of Example 2 and Example 1, which is located 8 times at 50 mm intervals, is made of aluminum near the surface. Example 3 positioned between foil and CFRP was further tested.

7A and 7B are graphs showing temperature distribution according to distance from the center of the heating element of the panel of the example and comparative example. As can be seen in Figure 7a, the temperature drop in Comparative Examples 2 and 3 is less than that in Comparative Example 1. This can be seen as an effect due to aluminum foil, which has high thermal conductivity. Additionally, it can be seen that the addition of cork board in Comparative Example 3 slightly increased the surface temperature compared to Comparative Example 2.

Figure 7b is a graph showing the temperature distribution according to the distance from the center of the heating element of the panel for Comparative Example 4 and Examples 1 to 3. In the case of Example 1 and Comparative Example 4, the heating element 30 located between the CFRP and the cork board forms a higher surface temperature and has a smaller temperature drop between the heating elements than when the heating element 30 is located in the second layer from the top among the CFRP layers. You can check that.

In the case of Example 2, in which the number of heating elements 30 was increased from Example 1, the surface temperature can be seen to have increased compared to Example 1, which is believed to be an increase in the maximum surface temperature due to the accumulated heat of adjacent heating films. Additionally, in the case of Example 3 in which the heating element 30 is located near the surface, the surface temperature can be seen to have increased compared to Example 1. Additionally, this can be confirmed through the steady-state heat distribution for Examples 2 and 3 in FIG. 8.

[Multiphysics anti-icing simulation under glaze ice conditions of the anti-icing and de-icing system of the aircraft wing]

Example

The present invention presents an experimental comparison of the thermal characteristics of the system through thermal finite element modeling (FEM) in a glaze ice condition in the anti-icing and de-icing system 1 of the aircraft wing according to an embodiment. Unlike the previous experiment conducted in a dry airflow, this corresponds to the simulation result in a wet airflow considering the freezing conditions of the dynamic airflow.

In addition, in the system using thermal finite element modeling (FEM), the airflow speed under flight conditions was set to 102 m/sec and the temperature was -6.65°C.

The embodiment is a system that forms 15 heating elements 30 (width: 25 mm) at 2 mm intervals on a 1 m long wing based on the laminated structure of Example 1 described above.

Under these conditions, the Angle of Attack (AOA) was set to 0° and 4°, and the heat flux was set to 2.5, 5.0, 7.5, and 10.0 kW/m2, respectively.

Figure 9a is a diagram showing airflow streamlines formed by 15 heating elements 30 at 2mm intervals and the results of icing simulation.

Referring to Figure 9a, it can be seen that a large amount of ice was formed regardless of the angle of attack when the system was not operating. And in the case of heat fluxes of 2.5 kW/m2 and 5.0 kW/m2, it can be seen that the ice was partially removed due to the heat generated from the heating element.

In particular, when the heat flux is 2.5kW/m2 and the angle of attack is 0°, run-back ice (a phenomenon in which water melted by the ice-making system flows backwards and refreezes in areas without an ice-making system) is found. Instead, no runback ice is formed at an angle of attack of 4°, but ice cones of considerable height are formed at low heat fluxes. This can be attributed to local flow acceleration on the upper surface.

On the other hand, when the heat flux is 7.5 kW/m2, it can be seen that all ice formed on the wing was removed regardless of the angle of attack.

In addition, Figure 10 shows the simulation results formed by 15 heating elements 30 at 2 mm intervals, (a) the airfoil cross-section temperature contour, (b) the surface temperature when the angle of attack is 0°, and (c) the surface temperature when the angle of attack is 4°. This is a drawing.

Referring to FIG. 10, an image of the surface temperature distribution of the aircraft wing according to the configuration of the embodiment is shown. In Figure 10(a), it can be seen that the surface temperature of the wing portion is uniform throughout both the upper and lower portions. In addition, in Figures 10(b) and 10(c), it can be seen that the surface temperature of the wing portion is distributed between 0 and 13°C, and is above 0°C in the entire area. Additionally, it can be seen that the temperature drop between heating elements in all areas is 2~3℃, so the possibility of forming runback ice is low.

Comparison example

The present invention presents an anti-icing and de-icing system (1) for aircraft wings according to a comparative example.

The comparative example is a system that forms five heating elements 30 (width: 25 mm) at 50 mm intervals on a 1 m long wing based on the laminated structure of Example 1 described above.

Other than this, the configuration was the same as that of the embodiment.

Figure 9b is a diagram showing airflow streamlines formed by five heating elements 30 at 50 mm intervals and the results of icing simulation.

Referring to Figure 9b, it can be seen that even though the heat flux was increased to 10.0 kW/m2, the ice was only partially removed from the wing portion. This can be seen to be caused by the backflow of partially melted ice and the gap between the heating elements 30.

That is, compared to the embodiment, as the distance between the heating elements 30 increases, the melted water moves to the rear by the system and then meets the cold surface temperature at the part where the heating element 30 is not attached and freezes again, forming runback ice. It can be seen that it was formed more easily.

In addition, Figure 11 shows the simulation results formed by five heating elements 30 at 50 mm intervals, (a) the airfoil cross-section temperature contour, (b) the surface temperature when the angle of attack is 0°, and (c) the surface temperature when the angle of attack is 4°. This is a drawing.

Referring to FIG. 11, an image of the surface temperature distribution of the aircraft wing according to the configuration of the embodiment is shown. In Figure 11(a), it can be seen that the temperature drop is large between the heating elements 30. In addition, it can be seen that the area where ice was formed is also the area where the temperature drop is large.

In particular, referring to FIGS. 11(b) and 11(c), even though the heat flux was increased to 10.0 kW/m2, the temperature drop between the portion formed at 0° or less between the heating elements 30 and the heating elements was 10° or more. You can see the change.

As a result, the configuration of reducing the gap between the heating elements 30 as much as possible and locating the heating element 30 as close to the surface of the wing as possible can reduce the temperature drop between the heating elements 30, thereby reducing the temperature drop between the heating elements 30. ·It can be confirmed that the anti-icing effect according to the heat flux is the most desirable in the ice-making system (1).

The anti-icing and de-icing system of the aircraft wing described above and shown in the drawings is only an example for carrying out the present invention, and should not be construed as limiting the technical idea of the present invention. Embodiments that have been improved and changed without departing from the gist of the present invention will be said to fall within the scope of protection of the present invention as long as they are obvious to those skilled in the art to which the present invention pertains.

1: Aircraft wing anti-icing and de-icing system
10: heat conductor layer
20: Carbon fiber stack layer
30: heating element
31: Electric heating element
32: High-performance plastic
40: insulation layer
50: Epoxy

Claims (13)

A carbon fiber stack layer in which several plies made of carbon fiber materials are laminated;
A heating element that supplies heat to remove ice that forms on the aircraft surface;
A heat conductor layer that conducts heat generated from the heating element; And
An anti-icing and de-icing system for an aircraft wing portion formed of an insulating material layer that prevents heat generated from the heating element from being conducted into the interior of the aircraft.
According to paragraph 1,
In the carbon fiber stack layer, the heat conductor layer is laminated on one side, and the insulation layer is laminated on the other side,
The insulation layer is in contact with the aircraft,
The heating element is characterized in that it is laminated at least one of between the heat conductor layer and the carbon fiber stack layer, between several layers of plies formed of the carbon fiber material, or between the carbon fiber stack layer and the insulation layer. Anti-icing and de-icing system for aircraft wings.
According to paragraph 2,
The heat conductor layer and the insulation layer are,
An anti-icing and de-icing system for an aircraft wing, characterized in that one side and the other side of the carbon fiber stack layer are respectively bonded to each other with epoxy.
According to paragraph 1,
The carbon fiber stack layer is,
An anti-icing and de-icing system for aircraft wings, characterized by multiple layers of plies formed of carbon fiber-reinforced polymer.
According to paragraph 1,
The heating element is,
An electric heating element,
An anti-icing and de-icing system for an aircraft wing, characterized in that it is made of heat-resistant, high-performance plastic formed to surround the exterior of the electrothermal heating element in order to insulate the electric current generated in the electrothermal heating element from the outside.
According to paragraph 1,
The heating element is,
Formed along the longitudinal direction of the aircraft wing,
An anti-icing and de-icing system for an aircraft wing, characterized in that a plurality of units are provided spaced apart at regular intervals in the width direction of the wing.
According to paragraph 1,
The heat conductor layer is,
An anti-icing and de-icing system for aircraft wings, characterized in that it is made of aluminum.
According to paragraph 1,
The insulation layer is,
An anti-icing and de-icing system for aircraft wings, characterized in that it is made of cork material.
Forming a carbon fiber stack layer by stacking several plies made of carbon fiber material;
Attaching a heating element that supplies heat to the carbon fiber stack layer to remove ice formed on the surface of the aircraft; And
A method of manufacturing an anti-icing and de-icing system for an aircraft wing formed by inserting and curing epoxy into the carbon fiber stack layer to which the heating element is attached.
According to clause 9,
A method of manufacturing an anti-icing and de-icing system for an aircraft wing, further comprising laminating a heat conductor layer that conducts heat generated from the heating element on one surface of the carbon fiber stack layer.
According to clause 9,
A method of manufacturing an anti-icing and de-icing system for an aircraft wing, further comprising laminating an insulation layer on the other surface of the carbon fiber stack layer to prevent heat generated from the heating element from being conducted into the interior of the aircraft.
According to claim 10 or 11,
Laminating at least one of the heat conductor layer and the insulation layer,
A method of manufacturing an anti-icing and de-icing system for an aircraft wing, further comprising bonding under vacuum pressure using epoxy.
According to clause 9,
In the step of attaching a heating element that supplies heat to the carbon fiber stack layer to remove ice formed on the surface of the aircraft,
The heating element is composed of an electrothermal heating element and heat-resistant High Performance Plastic formed to surround the outside of the electrothermal heating element to insulate the electric current generated in the electrothermal heating element from the outside,
A method of manufacturing an anti-icing and de-icing system for an aircraft wing, further comprising the step of plasma surface treatment of the surface of the high-performance plastic.

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