The present invention relates to turbomachinery, and more particularly to anti-icing and de-icing systems for aircraft engine nacelles, of which at least a portion may be fabricated from a composite material.
High-bypass turbofan engines are widely used for high performance aircraft that operate at subsonic speeds. As schematically represented in FIG. 1, a high-bypass turbofan engine 10 includes a large fan 12 placed at the front of the engine 10 to produce greater thrust and reduce specific fuel consumption. The fan 12 serves to compress incoming air 14, a portion of which flows into a core engine (gas turbine) 16 that includes a compressor section 18 containing low and high pressure compressor stages 18A and 18B to further compress the air, a combustion chamber 20 where fuel is mixed with the compressed air and combusted, and a turbine section 22 where a high pressure turbine 22A extracts energy from the combustion gases to drive the high pressure stages 18B of the compressor section 18 and a low pressure turbine 22B extracts energy from the combustion gases to drive the fan 12 and the low pressure stages 18A of the compressor section 18. A larger portion of the air that enters the fan 12 is bypassed to the rear of the engine 10 to generate additional engine thrust. The bypassed air passes through an annular-shaped bypass duct 24 that contains one or more rows of stator vanes, also called outlet guide vanes 28 (OGVs), located immediately aft of the fan 12 and its fan blades 26. The fan blades 26 are surrounded by a fan cowling or nacelle 30 that defines the inlet duct 32 to the turbofan engine 10 as well as a fan nozzle 34 for the bypassed air exiting the bypass duct 24.
The nacelle 30 is an important structural component whose design considerations include aerodynamic criteria as well as the ability to withstand foreign object damage (FOD). For these reasons, it is important to select appropriate constructions, materials and assembly methods when manufacturing the nacelle 30. Various materials and configurations have been considered, with metallic materials and particularly aluminum alloys being widely used. Composite materials have also been considered, such as graphite-reinforced epoxies, as they offer the advantage of significant weight reduction. However, in order to be meet aerodynamic and structural criteria, nacelles formed of composite materials encounter certain challenges. For example, laminar flow over wings, nacelles, and other surfaces is desirable to promote engine efficiency and improve specific fuel consumption (SFC). To achieve laminar flow on the nacelle, steps and gaps should be absent in its outer surface, from the inlet lip 36 to the maximum diameter 44 of the nacelle 30, or at minimum the length of the inlet outer barrel section immediately aft of the inlet lip 36. Though composites and their fabrication processes are well suited for producing single piece parts of this size with the required contour control and part weight, composite materials alone have not been capable of providing the impact resistance necessary to reliably survive in-flight bird strikes.
An additional issue concerning aircraft engine nacelles is that they are subject to icing conditions, particularly the nacelle leading edge at the inlet lip (36 of FIG. 1) while the engine is on the ground and especially under flight conditions. One well known approach to removing ice build-up (de-icing) and preventing ice build-up (anti-icing) on the nacelle inlet lip has been through the use of a hot air bleed system. An example is schematically represented in FIG. 1, in which engine-supplied bleed air flow is drawn from the compressor section 18 through piping 38 to the inlet lip 36, where the hot bleed air contacts the internal surface of the inlet lip 36 to heat the lip 36 and remove/prevent ice formation. The piping 38 includes a tube arrangement commonly referred to as a piccolo tube 40, which resides in an annular-shaped cavity of the nacelle 30 sometimes referred to as the D-duct 42. The tube 40 completely fills the D-duct 42 with the hot bleed air to ensure adequate heating of the inlet lip 36. While this type of system is effective, it requires a large amount of bleed air to fill the D-duct 42 and provide the thermal energy necessary to perform the anti-icing and de-icing functions. The hot air bled from the engine 10 results in a corresponding negative impact on engine performance and detracts from engine efficiency (SFC). Additionally, hot air bleed systems of the type represented can incur a significant weight penalty.
As an alternative, some smaller turbofans and turboprop aircraft engines have utilized electrical anti-icing systems, for example, resistance-type heater wires that may be attached to the interior surface of the inlet lip 36, or embedded in a boot bonded and/or mechanically attached to the interior surface of the inlet lip 36, or directly embedded in the inlet lip 36, such that heating of the lip 36 is through thermal conduction. However, such systems generally require excessive energy for de-icing and continuous anti-icing operation on large aircraft engines, such as high-bypass turbofan engines of the type represented in FIG. 1.