阅读说明：本技术 用于空气动力学系统的部件的防冰系统 (Anti-icing system for a component of an aerodynamic system ) 是由 S.塞尔瓦拉 S.桑加维尔 于 2019-07-11 设计创作，主要内容包括：公开了一种用于飞行器的空气动力学表面的防冰系统,表面具有流动面向侧和与该流动面向侧相对的向内面向侧,所述系统具有：穿孔板,其被构造成用于设置在表面中；加热源,其连接到该穿孔板；以及抽吸源,其被设置成通过穿孔板和加热源抽吸由加热源所融化的冰。(An anti-icing system for an aerodynamic surface of an aircraft is disclosed, the surface having a flow-facing side and an inwardly-facing side opposite the flow-facing side, the system having: a perforated plate configured for placement in a surface; a heating source connected to the perforated plate; and a suction source configured to suck the ice melted by the heating source through the perforated plate and the heating source.)

1. An anti-icing system for an aerodynamic surface of an aircraft, the surface having a flow-facing side and an inwardly-facing side opposite the flow-facing side, the system comprising:

a perforated plate configured for placement in the surface;

a heating source connected to the perforated plate; and

a suction source configured to suck the ice melted by the heating source through the perforated plate and the heating source.

2. The component of claim 1, wherein the heating source is integral with the perforated plate.

3. The component of claim 2, wherein the heating source comprises a drain hole extending between the perforated plate and the suction source.

4. The component of claim 3, wherein the anti-icing system comprises a honeycomb support structure on an inside of the perforated panel.

5. The component of claim 4, wherein the anti-icing system comprises a water collector on an inward facing side of the honeycomb support structure.

6. The component of claim 5, wherein the suction source is a pump fluidly connected to the sump.

7. The component of claim 6, wherein the anti-icing system comprises a rigid housing on an inward facing side of the water collector.

8. The component of claim 7, wherein the anti-icing system is coupled to a leading edge of the component.

9. The component of claim 8, wherein the component is a nacelle or an aircraft control surface on a wing and a tail.

10. The component of claim 8, wherein the component is a wing and tail.

11. A method of preventing icing with an anti-icing system located on a surface of a component of an aerodynamic system, the surface having a flow-facing side and an inwardly-facing side opposite the flow-facing side,

the method comprises the following steps:

heating the surface with a heating source of the anti-icing system, an

Providing suction with a suction source on an inward facing side of the heating source to draw water melted by heating of the surface by the heating source into the sump.

12. The method of claim 11, wherein the flow-facing side of the heating source comprises a perforated plate.

13. The method of claim 12, wherein the heating source comprises a drain hole extending between the perforated plate and the suction source.

14. The method of claim 13, wherein the anti-icing system comprises a honeycomb support structure on an inside of the perforated panel.

15. The method of claim 14, wherein the anti-icing system comprises a water collector on an inward facing side of the honeycomb support structure.

16. The method of claim 15, wherein the suction source is a pump fluidly connected to the sump.

17. The method of claim 16, wherein the anti-icing system comprises a rigid housing on an inward facing side of the sump.

18. The method of claim 17, wherein the anti-icing system is coupled to a leading edge of the component.

19. The method of claim 18, wherein the component is a nacelle.

20. The method of claim 18, wherein the component is an airfoil.

Background

Exemplary embodiments relate to the field of deicing components of aerodynamic systems, and more particularly to anti-icing systems for nacelle skins and control surfaces.

Safety is a central concern in the design of aircraft applications or in powered propulsion systems. The early stages of self-weight air flight have been a problem with icing on wings, propellers, engine air intakes, etc. Any ice build-up adds considerable weight and changes to the wing or inlet configuration, making it more difficult to fly and, in some cases, causing damage to the aircraft. In the case of jet aircraft, large pieces of ice flying off the leading edge of the engine inlet casing can severely damage the rotating fan and compressor blades extending through the airflow path and interact with the working fluid or other internal engine components and cause engine failure.

Many attempts have been made to overcome the problems and risks of icing of aircraft. Anti-icing systems for the inlet area of the nacelle of an aircraft propulsion system have been the focus of much research and development in the aircraft industry. For example, proposals have been made to mechanically vibrate the outer surface to loosen ice or to generate electromagnetic pulses in the skin of an aircraft to loosen ice. However, these systems tend to be heavy and complex, and only remove existing ice, rather than preventing icing.

It has been proposed many times to heat areas of aircraft prone to icing. Proposals for heating schemes range from microwave heating to the provision of hot gases through holes in the skin to resistive heating of the surface to actually burn the fuel adjacent to the surface prone to icing.

One of the most common anti-icing techniques is to direct hot gases into the housing adjacent to the potential ice formation areas. Typical patents describing such hot gas technology are U.S. patents nos. 3057154, 3925979, 3933327 and 4250250. In each case, the hot gas duct simply discharges hot gas into a casing, such as the leading edge of a jet engine casing or the leading edge of a wing. While generally useful, these systems are not entirely effective due to the complexity of the hot gas piping system.

In addition, with known techniques, the melted ice may flow down as a film of water, which may form an ice overflow (runback) downstream of the heater. For example, under the influence of aerodynamic forces, flooding may enter the engine core. In gas turbine engines, flooding can lead to core icing, component impact damage, and engine stall. On a wing, flooding ice may adversely affect aerodynamic profile, thereby affecting aerodynamic performance.

Disclosure of Invention

An anti-icing system for an aerodynamic surface of an aircraft is disclosed, the surface having a flow-facing side and an inwardly-facing side opposite the flow-facing side, the system comprising: a perforated plate configured for placement in a surface; a heating source connected to the perforated plate; and a suction source configured to suck the ice melted by the heating source through the perforated plate and the heating source.

In addition to, or as an alternative to, one or more of the above disclosed features and elements, the heating source is integral with the perforated plate.

In addition to, or as an alternative to, one or more of the above-disclosed features and elements, the heating source includes a drain hole extending between the perforated plate and the suction source.

In addition to or as an alternative to one or more of the above disclosed features and elements, the anti-icing system comprises a honeycomb support structure on the inside of the perforated plate.

In addition to one or more of the above disclosed features and elements, or as an alternative, the anti-icing system includes a water collector on an inward facing side of the honeycomb support structure.

In addition to, or as an alternative to, one or more of the above-disclosed features and elements, the suction source is a pump fluidly connected to the sump.

In addition to, or as an alternative to, one or more of the above-disclosed features and elements, the anti-icing system includes a rigid housing on an inwardly facing side of the water collector.

In addition to, or as an alternative to, one or more of the above-disclosed features and elements, an anti-icing system is coupled to a leading edge thereof.

In addition to or as an alternative to one or more of the above disclosed features and elements, the component is a nacelle or an aircraft control surface located on a wing or tail.

In addition to, or as an alternative to, one or more of the above disclosed features and elements, the component is a wing or tail.

Also disclosed is a method of preventing icing with an anti-icing system located on a surface of a component of an aerodynamic system, the surface having a flow-facing side and an inwardly-facing side opposite the flow-facing side, the method comprising: the surface is heated with a heating source of the anti-icing system and suction is provided at an inward facing side of the heating source with a suction source to draw water melted by heating the surface with the heating source into the sump.

Drawings

The following description should not be considered limiting in any way. Referring to the drawings wherein like elements are numbered alike:

FIG. 1 is a perspective view of an aircraft including an aerodynamic surface in which embodiments of the invention may be implemented;

FIGS. 2A and 2B illustrate one or more components of an aerodynamic system having an anti-icing system coupled thereto according to the present disclosure;

FIG. 3 is a view of an anti-icing system according to the present disclosure;

FIG. 4 is a view of a portion of an anti-icing system according to the present disclosure; and is

FIG. 5 is another view of an anti-icing system according to the present disclosure.

Detailed Description

A detailed description of one or more embodiments of the disclosed apparatus and methods is presented herein by way of illustration, not limitation, with reference to the figures.

FIG. 1 illustrates an example of a commercial aircraft 10 having aircraft engines surrounded (or otherwise carried) by a nacelle 20. The aircraft 10 includes two wings 22, each wing 22 may include one or more slats 24 and one or more flaps 26. The aircraft may also include ailerons 27, spoilers 28, horizontal stabilizer trim tabs 29, horizontal stabilizers 30 and rudders 31, and vertical stabilizers 32 (tail structures collectively referred to as empennages), each of which may be generally referred to as a "control surface" because they are movable under the aircraft's power system. The leading edges of the wings and of the nacelle are particularly susceptible to icing.

Turning now to fig. 2A, 2B and 3, a component 100 of an aerodynamic system is disclosed, which may be, for example, a nacelle of an aircraft engine 20 or a leading edge of a wing 20 or any control surface thereon. In general, component 100 may include a surface 110. The surface 110 may have a flow-facing side 120 and an inward-facing side 130 opposite the flow-facing side 120. Surface 110 may be coupled to anti-icing system 140. Anti-icing system 140 may, for example, be used to dynamically prevent and/or melt a layer of ice 135 that may accumulate on surface 110.

Anti-icing system 140 may include a heating source 150, the heating source 150 being flow-facing and disposed in a protected area of component 100, wherein surface 110 is porous to allow heat to transfer therethrough. The heating source may be electrically powered, but other heat sources may be used depending on the location of the anti-icing system 140. For example, heat may be obtained using air discharged from the engine, and may therefore be referred to as bleed air.

A suction source 160 may be disposed within the surface and draw fluid from the flow-facing side 120 to the inward-facing side 120. As shown in fig. 3, the heating source 150 is closer to the surface than the suction source 160. Suction source 160 may be a powered pump, but other suction sources may be used depending on the location of anti-icing system 140. For example, suction may be obtained through a conduit connected to a low pressure location in or around the engine.

In one embodiment shown in fig. 3, anti-icing system 140 located on the flow-facing side of heating source 150 may include a perforated plate 170. In one embodiment, heating source 150 is at least partially porous, allowing melted ice to travel between the flow side of heating source 150 and the inward facing side of heating source 150. In one embodiment, the anti-icing system may include drain holes 180 in the heating source 150. A drain hole 180 may extend between the perforated plate 170 and the suction source 160.

Turning to fig. 4 and 5, in one embodiment, anti-icing system 140 may include a honeycomb cell support structure 190. The honeycomb unit support structure 190 may support the heating source 150 and the perforated plate 170 on the flow-facing side thereof.

Anti-icing system 140 may include a water collector 200 located on an inward facing side of honeycomb cell support structure 190. The water collector may collect water formed from the layer of melted ice 135. The collected water may be used for cooling purposes, for example.

In one embodiment, anti-icing system 140 includes a rigid housing 210 located inside of water trap 200. Rigid housing 210 may be a rigid wall plate that fixedly supports anti-icing system 140.

Turning back to fig. 2, the component 100 may be a control surface of a nacelle or an aerodynamic system (such as a wing) to which an anti-icing system 140 is coupled. However, other components in an engine or aircraft that may experience icing due to exposure to airflow are also within the scope of the present disclosure. The surface 110 may be, for example, the outer skin at the leading edge of the nacelle 100, i.e. at the air intake lip 220, or at the leading edge of a wing, such as the outer skin on a slat.

The above disclosed embodiments may reduce/eliminate overflow of ice by draining the melted ice/water stream through porous heaters 150, for example, on the leading edge or one or more deicing zones of the nacelle 100. The disclosed embodiments may provide one or more drainage holes 180 at, for example, protected and unprotected areas of the nacelle 100 to remove a film of water formed by the melted ice 135. One or more drain holes 180 may be provided with the porous heater 150 and the external suction mechanism 160 may provide drainage for the flow of melted water. Suction source 160 may be activated along with heater 150 and may store and/or dispose of the drained water. Embodiments may reduce and/or eliminate ice spillage, improve aerodynamic performance, reduce drag, and improve fuel efficiency. In addition, embodiments may reduce downstream heating requirements (power and losses) due to reduced water and ice contact on the downstream surfaces of the engine.

The term "about" is intended to include the degree of error associated with measurement of a particular quantity based on equipment available at the time of filing this application.

The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the disclosure. As used herein, the singular forms "a", "an" and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms "comprises" and/or "comprising," when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.

While the disclosure has been described with reference to one or more exemplary embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the disclosure without departing from the essential scope thereof. Therefore, it is intended that the disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this disclosure, but that the disclosure will include all embodiments falling within the scope of the claims.