EP1055083B1

EP1055083B1 - Combustor flow controller

Info

Publication number: EP1055083B1
Application number: EP98963641A
Authority: EP
Inventors: John R. Tilston; John Austin
Original assignee: Qinetiq Ltd
Current assignee: Qinetiq Ltd
Priority date: 1997-12-17
Filing date: 1998-12-17
Publication date: 2002-11-06
Anticipated expiration: 2018-12-17
Also published as: EP1055083A1; DE69809295D1; DE69809295T2; AU1884199A; WO1999032827A1

Description

This invention relates to improved combustor arrangements for gas turbine engines and in particular is concerned with control of air flow to combustor zones.

Gas turbine engines include an air intake through which air is drawn and thereafter compressed by a compressor to enter a combustor at one or more ports. Fuel is injected into the combustion chamber by means of a fuel injector whence it is atomised, mixed with the compressed air from the various inlet ports and burnt. Exhaust gases are passed out of an exhaust nozzle via a turbine which drives the compressor. In addition to air flow into the combustion chamber through the air inlet ports, air also enters the combustion chamber via the fuel injector itself.

Conventional combustors take a variety of forms. They generally comprise a combustion chamber in which large quantities of fuel are burnt such that heat is released and the exhaust gases are expanded and accelerated to give a stream of uniformly heated gas. Generally the compressor supplies more air than is needed for complete combustion of the fuel and often the air is divided into two or more streams, one stream introduced at the front of the combustion chamber where it is mixed with fuel to initiate and support combustion along with the air in the fuel air mixture from the fuel injector, and one stream is used to dilute the hot combustion products to reduce their temperature to a value compatible with the working range of the turbine

Gas turbine engines for aircraft are required to operate over a wide range of conditions which involve differing ratios between the mass flows of the combustion and dilution air streams To ensure a high combustion efficiency, it is usual for the proportion of the total airflow supplied to the burning zone to be determined by the amount of fuel required to be burned to produce the necessary heat input to the to the turbine at the cruise condition. Often the chamber conditions are stoichiometric in that there is exactly enough fuel for the amount of air; surplus fuel is not completely burnt. However because of variability of the cycles and because air and fuel are never completely mixed there are always some oxides of nitrogen and unburnt fuel residues. An ideal air fuel mixture ratio at cruise usually leads to an over rich mixture in the burning zone at high power conditions (such as take-off) with resultant unburnt hydrocarbon and smoke emission. It is possible to reduce smoke emission at take-off by weakening the burning zone mixture strength but this involves an increase in primary zone air velocity which makes ignition of the engine difficult to achieve, especially at altitude.

The temperature rise of the air in the combustor will depend on the amount of fuel burnt. Since the gas temperature required at the turbine varies according to the operating condition, the combustor must be capable of maintaining sufficient bum over a range of operating conditions. Unwanted emissions rise exponentially with increase in temperature and therefore it is desirable to keep the temperature low. With increasingly stringent legislation against emissions, engine temperature is an increasingly important factor, and operating the combustor at temperatures of less than 2100K becomes necessary. However at low temperatures, the efficiency of the overall cycle is reduced.

It is a requirement that commercial airliners can decelerate rapidly in the case of potential collision. In order to decelerate a gas turbine from high power to low power, the fuel flow to the engine is reduced. Although the reduction in fuel flow is almost instantaneous, the rate of reduction of engine airflow is relatively slow because of the inertia of rotating parts such as turbines, compressors, shafts etc. This produces a weak mixture of fuel and this increases the risk of flame extinction. It is not always easy to relight the flame especially when the combustor is set to run weakly and at high altitude. Because modern combustors invariably operate in lean burn principles in order to reduce oxide of nitrogen emissions, combustors need to be operated as close to the lean extinction limit at all engine operating conditions. If margins are set wide enough to prevent flame extinction then emissions performance is compromised.

Combustion is initiated and stabilises in the pilot zone, the most upstream section of the combustor. Low power stability requires rich areas within the primary zone of the combustor, enabling combustion to be sustained when the overall air/fuel ratio is much weaker than the flammability limit of kerosene. In traditional combustion systems rich regions can occur in the combustor due to poor mixing and poor atomisation resulting in large droplets of fuel being formed.

Conventional gas turbine engines are thus designed as a compromise rather than being optimised, because of consideration of the above mentioned conflicting requirements at different operating conditions. New "staged" design of combustors overcome the problems to a limited extent. These comprise two combustion zones, a pilot zone and a main zone, each having a separate fuel supply. Essentially this type of combustor is designed such that a fixed flow of about 70% enters the combustor at the main zone and about 30% of the air flows to the pilot zone. In such systems the air/fuel ratio is determined by selecting the amount of fuel in each stage. The air/fuel ratio governs the temperature which determines the amount of emissions. Current gas turbine engine trends are towards increased thrust/weight ratios which require the engine to perform at higher operating compression ratios and wider ranges of combustor air/fuel ratios. Future gas turbine combustion systems will be expected to perform at higher inlet temperatures and richer air/fuel ratios. Because there is little variability in the airflow proportions to the main stage and pilot stage the amount of optimisation achievable for each operating condition is reduced. Even these combustor designs will suffer from either high nitrogen oxide and smoke emissions at full power, or poor stability at low power.

It is therefore desirable to improve control of the amount of fuel, air and air/fuel ratio in each combustor zone to reduce the problems of weak flame extinction, emissions of oxides of nitrogen and unburnt fuel at all operating conditions, whilst maintaining good efficiency and performance.

Conventionally, as shown in GB 785,210, this can be achieved by diverting a main airflow flowing through a main conduit into one of two subsidiary conduits by injecting under pressure into the main airflow a controlling air stream. Such devices which use a valve system to inject, under pressure, a controlling air stream when desired are described in US-A-3,631,675, US-A-3,910,035 and DE-A-2657707. However, this requires a separate compressor which is disadvantageous in terms of cost and weight. Alternatively, GB 1,184,683 discloses a system whereby a suction action is utilised. However, this is achieved by bleeding compressed air out of the engine resulting in a loss of engine efficiency.

It is an objection of the invention to provide enhanced means by which air flow can be controlled.

According to a first aspect of the present invention, a flow controller for supplying air to a combustor comprises a conduit and a control port, the conduit including a main section dividing into at least two secondary sections at a junction and the control port being positioned in the main section adjacent to the junction, characterised in that the control port is connected to a reservoir; and wherein, in use, a change in the flow rate of a main airflow flowing through the main section of conduit causes a control airflow to flow either in to or out of the control port whereby the main airflow is selectively diverted into one or other of the secondary sections of conduit.

A change in the flow rate of a main airflow results in a change in the static pressure of the main airflow which produces a pressure differential between the conduit adjacent to the port and the reservoir. The pressure differential causes the control airflow until pressure equalisation, the duration of the flow depending, amongst other things, on the size of the reservoir.

The control airflow flowing either in to or out of the control port causes a main airflow flowing through the main section of conduit to coanda around a surface of the main section whereby the main airflow is selectively diverted into one or other of the secondary sections of conduit. Ideally, the flow controller comprises at least one arcuate surface common to both the main section and a secondary section.

A skilled person would interpret coanda in relation to the coanda effect, the coanda effect being the tendency of a fluid jet to attach to a downstream surface roughly parallel to the jet axis. If this surface curves away from the jet the attached flow will follow it deflecting from the original direction (Dictionary of Science and Technology, Larousse 1995).

Preferably, the control port is connected to the conduit further upstream of the junction so as to form a control loop.

Preferably, the main section of conduit comprises a convergent-divergent duct; wherein, in use, the control airflow flowing either in to or out of the control port is caused by pressure differential across the duct.

According to a second aspect of the present invention, a gas turbine combustor comprises a flow controller as described above. Ideally, the flow controller comprises two secondary sections of conduit connected to two different zones within the combustor. In a preferred embodiment, the flow controller comprises one secondary section of conduit connected to a pilot combustion zone within the combustor and another secondary section of conduit connected to a main combustion zone.

In this way the proportion of flow to the main combustor zone and the pilot zone can be selectively altered without mechanical means. This provides robust control of flow with high reliability.

A combustor incorporating a flow controller according to the present invention will now be described, by way of example only, with reference to the drawings of which:

Figure 1 shows a schematic sectional view of a combustor incorporating a flow controller of the present invention;

Figure 2 shows the combustor of figure 1 in greater detail;

Figures 3a to d show the operation of the flow controller of the combustor of Figure 1 at various operating conditions.

Figure 4 shows alternative embodiments of the flow controller comprising one or more control ports in various locations.

Figure 1 shows a schematic view of a combustor incorporating a flow controller of the present invention. The combustor 1 comprises a main (high power) combustor zone 2 and pilot (low power) 3 combustor zone. Attached to the pilot zone is a primary fuel injector 4. Air flow into the combustor enters through a common entry point and a flow controller 5 which subdivides into two conduits one, 6, which leads to the main zone and the other, 7 to the pilot zone.

Figure 2 shows the flow controller for the combustor in more detail. The figure also shows a series of planes P1 to P4, in order to assist in the description of the flow controller. The air supply to the combustor is from a flow controller which comprises a main conduit 8 which divides into two separate sub conduits at P3, of which one (6) enters the main combustion zone, and the other (7) enters the pilot combustion zone. Upstream of the divergence formed by the subdivision of the conduit is located a control port 9. Port 9 is connected to a reservoir 10 which includes a valve 11 located on the other side which connects to the same pressure as at P1. A pressure difference exists from P1 to P4 such that air flows from P1 to P4. The conduit from P1 to P3 acts as a venturi. From P1 to P2 the flow cross section is such that flow of air accelerates and the static pressure falls to P2 which is lower than P1. This ensures that when valve 11 is open air will flow into the device from the control loop 16 and the control port. Downstream of P2 is a diffuser.

The angle of the diffuser is sufficiently large such that flow will coanda or attach to one or other of the outer walls. Some degree of diffusion and pressure recovery will take place and is essential in order for flow acceleration and pressure reduction at plane 2.

The operation of the embodiment described above will now be described with reference to Figure 3. Figure 3a shows the operation at idle condition. The reservoir pressure is neutral and the valve is opened such that control flow is injected through control port into the main flow where it acts as a boundary layer trip such that the main flow separates from wall to wall. The air flow now flows through sub conduit 6 to the main zone of the combustor. Fig. 3b shows that on acceleration, main flow is switched back to the sub conduit which leads to the pilot zone of the combustor by shutting valve 11. Control flow is sucked into the control port because the reservoir pressure is low. Figure 3c shows that at cruise condition the valve remains shut and the reservoir pressure is neutral. Air continues to flow to the pilot zone. On deceleration (Fig 3d) the reservoir pressure is overpressurised and flow out of the control port causes the main flow to divert into the conduit to the main zone.

The above described embodiment describes how control flow through a port in the flow controller can selectively divert flow, and flow control of air to each combustor zone is automatically selected.

In the embodiment only one control port is described. However any number of control ports in the vicinity of the divergence will have a controlling effect to direct the main air flow. Figure 4 shows four possible locations of control ports. Over-pressure (flow into conduit) at any of ports 12 to 14 will tend to divert flow to the sub-conduit 7 and conversely underpressure at any of ports 13 of 15 will tend to divert the flow to this sub-conduit.

Overpressure at any of ports (flow to main conduit) 13 or 15 will divert flow to the sub-conduit 6 and conversely under-pressure in any of

ports

12 or 14 will tend to divert the flow to the sub-conduit 6.

The diverted flow is stable in either of the two states even if there is no applied control flow. However the control flow is preferably provided by selective over(or under-) pressure at one of two

ports

12, 13 oppositely located adjacent the respective sub-conduit.

Claims

A flow controller (5) for supplying air to a combustor comprising a conduit (6,7,8) and a control port (9), the conduit including a main section (8) dividing into at least two secondary sections (6,7) at a junction and the control port being positioned in the main section adjacent to the junction characterised in that the control port is connected to a reservoir (10); and wherein, in use, a change in the flow rate of a main airflow through the main section of the conduit causes a control airflow to flow either in to or out of the control port whereby the main airflow is selectively diverted into one or other of the secondary sections of the conduit.
A flow controller according to claim 1 wherein the control port is connected to the conduit further upstream of the junction so as to form a control loop.
A flow controller according to claim 2 wherein the main section of conduit (8) comprises a convergent-divergent duct; wherein, in use, the control airflow flowing either in to or out of the control port is caused by a pressure differential across the duct.
A gas turbine combustor (1) comprising a flow controller (5) according to any preceding claim.
A gas turbine combustor (1) according to claim 4 wherein the flow controller (5) comprises two secondary sections of conduit (6,7) connected to two different zones within the combustor.
A gas turbine combustor (1) according to claim 4 wherein the flow controller (5) comprises one secondary section of conduit (7) connected to a pilot combustion zone within the combustor and another secondary section of conduit (6) connected to a main combustion zone within the combustor.