During operation of a gas turbine engine installed as a prime mover in an aircraft, heat is generated in various parts of the engine, its accessories and its associated fluid-flow systems. Thus, heat is obviously generated during combustion of the fuel in the engine's combustion system. Though a very high percentage of this exits the engine by way cf the hot exhaust jet, or is turned into work within the engine, some heat from the combustion process is absorbed by engine components and systems. Heat is also generated by frictional effects and similarly input to engine components and systems. These frictional effects may be mechanical, such as exist between relatively moving parts of the engine or its accessories; aerodynamic, due to drag and stagnation effects in the engine's gas passages; or fluid, due to churning of oil in the lubricating oil system, fuel pumping losses, pipe flow losses, etc.
The heat input to the installation is to a large extent dissipated from it by processes of conduction, convection and radiation, but heat is also transported around the installation by the engine's fluid flow systems. These can be used in order to prevent local overheating, to raise the efficiency of heat usage, and to facilitate desirable heat exchange either internally of the installation, or by way of dumping excess heat to the environment.
It has therefore become common practice in aircraft gas turbine installations to transfer heat from the oil in the engine's lubrication oil system to the fuel in the engine's fuel system by means of a heat exchanger designated a "fuel cooled oil cooler" (FCOC) or similar. The lubricating oil of course picks up a lot of the heat generated in the engine's bearings, in the oil pumping process, and from other sources, and this is transferred to the fuel in order both to prevent the oil overheating and to improve specific fuel consumption by raising the fuel temperature prior to combustion.
Another known practice is to provide fuel and/or oil systems with air-cooled heat exchangers for dumping excess heat from these systems into the atmosphere. In turbofan engines such heat exchangers are sited within the bypass duct so that the heat can be efficiently passed to the fan air stream. However, such heat exchangers cause a loss of thrust in the fan air stream and impose a drag penalty, leading to higher fuel consumption.
Under certain operational conditions--for example, when fuel held in an aircraft's fuel tanks is too warm due to heat soaking of the aircraft or prior fuel storage at high ambient ground temperatures, or when short flight times allow inadequate time at stratospheric altitudes for the cooling effect of the cold air on the aircraft structure to keep fuel tank temperatures low--heat input from the lubricating oil to the fuel can cause the fuel temperature before combustion to become too high for safety due to the danger of vaporisation in the fuel system, e.g. excessive cavitation during pumping.
On the other hand, there are many circumstances when fuel tank temperature is low, even though temperatures in the installation's fluid flow systems are high, and in this case rejection of heat to the environment is wasteful, having an adverse effect on the engine's specific fuel consumption. Furthermore, sub-zero fuel tank temperatures can lead to icing problems under some atmospheric conditions and warming of the fuel in the tanks to above 0.degree. C. helps to avoid such problems.
One object of the present invention is therefore to provide a convenient means of managing the heat flows in the installation's fluid flow systems in such a way that excessively high or low fuel and oil temperatures can be avoided and the heat capacity of the fuel in the fuel tanks utilised to help in this.
The problems are exacerbated by the continued advance of aircraft gas turbine technology in terms of reduction of fuel consumption by such means as increased compressor compression ratios, higher combustor exit temperatures and increased rotor speeds. Combined, these factors present considerable challenges in efficiently managing the heat generated within the relatively small volume of the engine. It is thus an object of the present invention to provide an improved capacity to manage large quantities of heat in aircraft gas turbine installations.
Besides the lubricating oil system of an engine, another closed-circuit oil system associated with large engines is that used for lubricating and cooling an electrical generator which is driven by the engine for providing electrical power to engine and aircraft systems. Unlike the engine's oil system, the generator's oil system is conventionally considered as separate from the rest of the engine installation for heat exchange purposes, and is normally provided with its own ACOC if necessary to avoid placing any additional load on the heat management capacity of the other fluid-flow systems, even though an ACOC imposes a drag penalty due to its position in an airstream flowing through or past the engine and even though under many engine operational conditions an additional heat input to the fuel would benefit efficiency. Consequently, it is a desirable object of the present invention to facilitate efficient integration of such generator oil systems with other fluid-flow systems of the engine installation in terms of their heat-exchange relationships, thereby providing the installation with a more comprehensive heat management capability and avoiding the need to give the generator's oil system its own ACOC.