Combustion equipment for gas turbine engines

Combustion equipment designed to reduce smoke levels and the production of nitrogen oxides has a fuel burner mounted in the upstream wall of a combustion chamber, the fuel burner comprising a central duct partially surrounded by an annular duct, each of the ducts having an array of swirl vanes at their upstream ends and fuel inlet apertures downstream of the respective arrays of swirl vanes. The fuel supply to each of the ducts can be controlled to apportion the fuel flow between the two ducts in dependence of an engine parameter, such as speed, so that at low speeds a majority of the fuel is injected into the annular duct while at higher engine speeds, a majority of the fuel is injected into the central duct.

This invention relates to combustion equipment for gas turbine engines. 
BACKGROUND OF THE INVENTION 
Combustion equipment design has been changed over recent years from the 
type using a fuel burner employing the fuel pressure jet principle to a 
fuel burner using the air-assisted principle. The primary motivation for 
this change has been the requirement to reduce the production of smoke as 
the pressure level within gas turbine high pressure spools has increased. 
Usually air-assisted burners feature the injection of fuel tangentially 
into a circular or annular air passage in which there is a high velocity 
air flow. This creates a cylindrical liquid sheet adjacent to the wall of 
the air passage and the resulting fuel placement in the combustion chamber 
of a gas turbine engine is usually in the form of a hollow cone. The 
fuel/air mixture is thus consequently very rich about the fuel sheet, and 
large amounts of smoke can still be produced. At low engine power 
conditions, the spray can have a wide range of droplet sizes which are 
related to the thickness of the fuel sheet presented to the incident 
airstream. 
It is one of the objects of the present invention therefore to provide 
combustion equipment for a gas turbine engine having a fuel burner which 
will provide a uniform spray of finely atomized fuel. A further object of 
the present invention is to provide combustion equipment which will 
produce reduced quantities of objectionable exhaust emissions, such as 
nitrogen oxides. 
The formation of nitrogen oxides is dependent upon a number of 
inter-related factors, including the temperature of combustion (the higher 
the temperature, the more nitrogen oxides are produced), the 
concentrations of nitrogen and oxygen in the fuel/air mixture, and the 
residence time of the combustion products in the combustion chamber. In 
the case of the residence time, low nitrogen oxides emissions can be 
achieved by having a short residence time with efficient combustion or by 
having a longer residence time with less efficient combustion so that the 
temperature is maintained at a low value and is insufficient for 
significant quantities of nitrogen oxides to be formed. 
Over the normal operating range of a gas turbine engine the conditions vary 
considerably in the combustion equipment because of variations in the air 
and fuel flow rates, and changing pressures and temperatures, and thus it 
is very difficult to reduce the formation of nitrogen oxides at all engine 
speeds. 
In our prior British Pat. No. 1,427,146 a tubular primary intake containing 
a fuel injector is provided in the upstream wall of the flame tube. An end 
cap is located at the downstream end of the tubular intake, to define an 
annular radially directed gap between it and the end of the tubular 
intake. This gap directs the fuel/air mixture radially into the flame tube 
creating a first toroidal vortex substantially upstream of the gap, and a 
second toroidal vortex of opposite hand substantially downstream of the 
gap. This arrangement has the ability to achieve high combustion 
efficiencies at ground idling engine speeds without detriment to high 
speed performance. 
BRIEF DESCRIPTION OF THE INVENTION 
According to the present invention combustion equipment for a gas turbine 
engine comprises a fuel burner comprising a hollow duct intended to 
receive a flow of air, first swirl means located adjacent to the upstream 
end of the hollow duct, an annular outer duct at least partially 
surrounding the hollow duct, second swirl means located adjacent to the 
upstream end of the annular outer duct and means for injecting fuel into 
each duct downstream of the first and second swirl means. 
The fuel may be injected into each duct normal to the axis of the ducts or 
at an acute angle to the axis of the ducts. 
The fuel may be injected into each duct from the outer wall of the hollow 
duct, the fuel being injected radially inwardly into the hollow duct and 
radially outwardly into the annular duct. 
Alternatively or additionally a fuel injector may be provided at the centre 
of the hollow duct. Alternatively additional means may be provided for 
injecting fuel into the outer duct from the outer wall thereof. 
Preferably the means for injecting fuel into the ducts is controlled such 
that the fuel is injected into only one or both of the ducts at a time in 
dependence upon various engine parameters, such as engine speed and power 
requirements. 
When the fuel is injected into both of the ducts the ratio of the volumes 
of fuel injected into each duct may be varied in dependence upon various 
engine parameters. 
Thus at low engine power the majority or all of the fuel may be injected 
into the outer duct and at high engine power the majority or all of the 
fuel may be injected into the hollow duct. 
Preferably the downstream end wall of the hollow duct is flared whereby a 
radial component of direction is imported to the fuel/air mixture issuing 
from the hollow duct. The downstream end wall of the annular outer duct 
may also be flared to import a radial component of direction to the 
fuel/air mixture issuing from the outer duct. 
The invention also comprises a gas turbine engine having combustion 
equipment as set forth above.

DETAILED DESCRIPTION OF THE INVENTION 
With reference to FIG. 1, a gas turbine engine generally indicated at 10 
comprises, in axial flow series, an air intake 11, compressor means 12, 
combustion apparatus 13, turbine means 14 and an exhaust nozzle 15. 
Combustion apparatus 13 comprises a plurality of separate, substantially 
cylindrical combustion chambers, one of which can be seen at 16, 
circumferentially mounted around the axis of the engine 10 in the manner 
which is generally described as a "cannular" array. 
Each combustion chamber consists of an annular wall 18 and an upstream end 
wall or base plate 20. Both the wall 18 and the base plate 20 are provided 
with small holes or orifices 22 to permit air to enter the combustion 
chamber for wall cooling purposes and larger holes 24 are provided in the 
wall 18 to permit combustion air to enter the chamber. The wall 18 is also 
provided with cooling air or dilution air holes 26, which air cools the 
combustion gases to a temperature acceptable to the turbine blades located 
downstream of the combustion chamber 16. 
Mounted in the centre of the base plate 20 is a fuel burner 28 which 
consists basically of two coaxial tubes 30 and 32, the outer tube 32 
surrounding the inner tube 30, and being slightly shorter than the inner 
tube 30 to define an annular passage 34 between the tubes. At the upstream 
end of the inner tube 30 is located a set of swirl vanes 36, and at the 
upstream end of the outer tube 32 is located a further set of swirl vanes 
38, these vanes also serving to support the inner tube 30 in position. The 
two sets of swirl vanes 36, 38 can be arranged to generate swirling flows 
of either the same hand or of opposite hand. 
The inner tube 30 is provided at its upstream end with two annular fuel 
manifolds 40 and 42 and holes 44 connect the manifold 40 with the interior 
of the tube 30, and holes 46 connect the manifold 42 with the passage 34. 
The holes 44 are arranged substantially perpendicularly to the axis of the 
burner 28. 
The downstream ends of each of the inner tube 30 and the outer tube 32 are 
flared outwardly, the outer tube 32 terminating slightly upstream of the 
inner tube 30 to provide a substantially radially facing annular gap 47 at 
the end of the passage 34. 
The supply of fuel to the two manifolds 40, 42 is controlled by a fuel 
scheduler 50 which receives fuel from a supply 52 and apportions the fuel 
to the manifold in dependence of an engine parameter 54, such as engine 
speed, compressor delivery pressure etc., as described in more detail 
below. 
During operation of the burner, air enters the passage 34 and the interior 
of the tube 30 through the swirler vanes 38 and 36 respectively which 
impart a high degree of swirl to the air. The air passes out of the gap 47 
and the end of the tube 30 in the direction of the arrows to create a 
first toroidal vortex 100 substantially upstream of the gap 46 and a 
second toroidal vortex 200 substantially downstream of the gap 46. These 
vortices are assisted by the incoming air through the holes 22 in the base 
plate 20 and the air entering the combustion chamber through the holes 24. 
At low engine power the fuel scheduler 50 apportions the majority or all of 
the fuel to the manifold 42 from whence it is injected into the outer duct 
34 and thence into the first toroidal vortex 100. By design of the 
combustion chamber, the equivalence ratios [Fuel/air ratio 
(actual)/Fuel/air ratio (stoichiometric)] of the first and second vortices 
can be optimized for maximum combustion efficiency at idle engine speeds. 
Since the fuel jets issuing from the holes 46 are angled to be 
substantially perpendicular to the swirled airflow, the very high relative 
velocity between the air and fuel achieves maximum atomisation. 
As fuel flow/power is increased the fuel scheduler apportions a greater 
proportion of the fuel to the manifold 46 from whence it is injected into 
the inner tube 30 and thence directly into the second vortex 200. Thus at 
high power, because only a portion of the fuel is directed into the first 
vortex 100 its equivalence ratio is maintained below that which results in 
excessive smoke production. 
The equivalence ratio of the second vortex 200 at high power is dictated to 
a large extent by airflow proportioning necessary to give optimum carbon 
monoxide consumption at idling speeds, but generally that fuel/air ratio 
in the second vortex 200 is similar to that of a conventional fuel burner. 
However, as the inner tube 30 fuel is intended to be pre-aerated, smoke 
production in the second vortex is minimal. 
The differential fuelling of the two vortices ensures that the first vortex 
remains relatively rich in fuel so reducing the possibility of the 
production of nitrogen oxides. 
At full power the two vortices can have equivalence ratios too rich for the 
production of nitrogen oxides followed by rapid dilution by air from the 
holes 26 to a low equivalence ratio of, for example, below 0.7. 
The combustion equipment therefore offers a great control over the local 
equivalence ratios within the combustion chamber thus enabling nitrogen 
oxide and smoke production to be maintained at low levels at different 
engine powers. 
The combustion equipment can be used not only for a "cannular" array, but 
also for tubo-annular combustion chamber or an annular combustion chamber.