Patent Application
20060032983
The present invention relates to de-icing and anti-icing systems for use with nacelles housing aircraft engines.
Turbofan nacelles typically require inlet de-icing for safety reasons. Prior art engine inlet anti-icing systems commonly employ a thermal source, such as hot air bled from the engine core or an electrical heating element, for providing heat to remove ice build-up on the inlet surfaces. However, using bleed from the engine core reduces overall engine efficiency and electrical systems draw electrical power which impose a non-propulsive load on the engine. Opportunities for improvement therefore exist.
It is therefore an aim of the present invention to provide an improved anti-icing system for an aircraft engine nacelle.
Therefore, in accordance with the present invention, there is provided a nacelle for housing a gas turbine engine, the nacelle comprising: an inlet lip defining a leading edge of the nacelle; a conduit located within the inlet lip, the conduit having a fluid circulating therein, the fluid providing a heat source; and an energy attenuating member located within the inlet lip and disposed between the leading edge and the conduit, the energy attenuating member providing protection to the conduit from foreign object damage and being thermally conductive such that heat transfer communication between the conduit and the leading edge is provided.
Also in accordance with the present invention, there is provided a system for preventing ice build up on an inlet lip of a nacelle housing a gas turbine engine, the system comprising: a cavity extending within the inlet lip and partly defined by a leading edge thereof; first means for providing a fluid circulation within the cavity; a hot fluid circulating within the first means; and second means for providing heat transfer communication between the first means and the leading edge and for providing foreign damage protection to the first means, the second means filling a free space between the first means and the leading edge.
Further in accordance with the present invention, there is provided a method of producing a foreign object damage tolerant anti-icing system for an inlet lip of a nacelle housing a gas turbine engine, the method comprising the steps of: defining a circumferential passage within the inlet lip; defining a free space between the circumferential passage and a leading edge of the inlet lip; filling the free space with a material, the material having sufficient thermal conductivity to enable heat transfer communication between the circumferential passage and the leading edge and having sufficient impact energy resilience to protect the circumferential passage from foreign object damage; and connecting the circumferential passage to a system circulating a hot fluid, the hot fluid circulating through the circumferential passage being adapted to provide heat to the inlet lip through the circumferential passage and the material.
There is further provided, in accordance with the present invention, a method of preventing foreign object damage to an anti-icing system in an inlet of a nacelle for housing a gas turbine engine, the method comprising: providing a conduit defining a fluid passage within the inlet lip; defining a space between the conduit and a leading edge of the inlet lip, and disposing an energy attenuating member within the space; enabling heat transfer communication between the conduit and an outer surface of the inlet lip via the energy attenuating member; and circulating a hot fluid within the conduit.
Still other aspects of these and other inventions will become apparent upon review of the description below.
Reference will now be made to the accompanying drawings, showing by way of illustration preferred embodiments of the present invention in which:
Referring to
The nacelle 10 is generally tubular, having an outer surface 31 and an inner surface 33 substantially parallel to one another and radially spaced apart to define a hollow cavity 29 therebetween. The circumferential inner surface 33 of the nacelle 10 defines the air flow passage to the engine at the upstream end thereof, and defines the annular bypass airflow passage 18 further downstream. At the most upstream end of the nacelle 10 is disposed an inlet lip 28. Within the annular hollow cavity 29 at the inlet lip 28 of the nacelle 10 is disposed a combined anti-icing and oil cooling system 30. A combined anti-icing system and oil cooler is disclosed in the applicant's co-pending application U.S. Ser. No. 10/628,368 filed Jul. 29, 2003, the contents of which is incorporated herein by reference. While efficient, the disposition of the system is such that it could be susceptible to foreign object damage. Should such damage occur, substantial repair costs and engine and/or aircraft down time may result. A more damage tolerant system is therefore desired, and will now be described.
Referring to
The upstream portion of the hollow cavity 29 within the inlet lip 28 includes an energy attenuating member 86, which has a high thermal conductivity such that heat transfer communication is maintained between the tube 34 and the outer surface the inlet lip. The energy attenuating member 86 is disposed between the tube 34 and the leading edge of the inlet lip, and preferably at least partially surrounds the tube 34. The energy attenuating member 86 preferably comprises a high thermal conductivity graphite foam, having a thermal conductivity similar to solid aluminum. The energy attenuating member 86 is such as to offer appropriate impact energy resilience against foreign object damage. Thus, the energy attenuating member 86 will crumple when impacted by a large foreign object striking the inlet lip 28, thereby dissipating the energy of the foreign object strike without significantly damaging the tube 34. Upon smaller foreign object damage strikes, the inlet lip of the nacelle may only be cracked or punctured by the object, and the energy attenuating member will tend to restrain the object such that normal operation of the anti-icing system will not be affected. The thermal conductivity properties of the energy attenuating member 86 allows heat transfer communication between the wall of the tube 34 and the annular sheet metal lip 36, as well as between the wall of the tube 34 and the inner and outer surfaces 33 and 31 of the nacelle 10, such that heat transfer by conduction can occur therebetween.
Hot engine oil having cooled the turbofan engine 14 is thus circulated through the oil passage 40, preferably continuously, before it is returned to the engine. Accordingly, heat transfer communication between the hot engine oil flowing through the oil passage 40 and the inlet lip icing regions of the nacelle inlet lip 28, through the high thermal conductivity material 86, allows heat from the hot engine oil to be transferred to an outer surface 32 of the inlet lip 28, thereby melting any ice formed thereon and keeping the outer surface 32 sufficiently warm in order to prevent any ice build-up, while simultaneously cooling the engine oil.
The system 30 as described thus allows the simultaneous cooling of the engine oil and de-icing of the inlet lip 28. In addition, the material 86 filling the inlet lip 28 provides foreign object damage protection to the tube 34. A small foreign object which punctures the outer surface 32 of the inlet lip 28 will likely be retained by the material 86 and as such will not interfere with the normal operation of the system 30. The material 86 will exhibit local damage only, which is easier and less costly to repair than damage to the tube 34.
A control system is provided for managing the anti-icing system to ensure that the necessary heat transfer and engine oil circulation is maintained.
In an alternate embodiment, a heat transfer fluid other than the engine oil is circulated through the passage 40, such that the tube 34 is the condenser component of a thermosyphon loop heated by a hot coil. The heat transfer fluid thus circulates through the passage 40 partly in a gaseous or vaporized form such as to be condensed therein. The heat transfer fluid possesses suitable vapor pressure characteristics, is non flammable, and is compatible with the materials it comes in contact with such as the material forming the tube 34. The heat transfer fluid also preferably has a zero ozone depletion potential (ODP) and a low global warming potential (GWP). However in this case, a conventional oil cooler is separately provided in the gas turbine engine for cooling the engine oil.
The embodiments of the invention described above are intended to be exemplary. Those skilled in the art will therefore appreciate that the forgoing description is illustrative only, and that various alternatives and modifications can be devised without departing from the spirit of the present invention. Accordingly, the present is intended to embrace all such alternatives, modifications and variances which fall within the scope of the appended claims.
