GB2073396A

GB2073396A - Gas turbine combustion chambers

Info

Publication number: GB2073396A
Application number: GB8010674A
Authority: GB
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1980-03-29
Filing date: 1980-03-29
Publication date: 1981-10-14
Also published as: GB2073396B

Abstract

A combustion chamber includes cooling air inlets 40, 42, 44 of the V-ring type or similar in which cooling air is introduced along the walls in both the upstream and downstream directions and the combustor flow pattern encourages this cooling air flow after it has performed the cooling function, to become entrained with the combustion and/or dilution phases. By this method the layer of cool air at the combustor exit is relatively thin or practically non-existent and the cooling air ceases to be parasitic and can be used in the burning and dilution phases. <IMAGE>

Description

SPECIFICATION Improvements in or relating to gas turbine engine combustion chambers This invention relates to gas turbine engine combustion chambers and is concerned with the cooling air used to cool the walls of such combustion chambers.

The conventional method of cooling the walls of gas turbine engine combustion chambrs is to introduce air from the engine compressor in such a way that a stream of relatively cool air passes down the walls separating as well as possible, the walls from the high temperature combustion gases. This method is known as film cooling. At discrete axial distances along the walls, the cooling film is refreshed and as a result, quite a thick layer of relatively cool air is built up at the exit of the combustion chamber.

This method of cooling has a number of disadvantages. The combustion products and unburnt fuel pass into the stream of cooling air which greatly reduces the further reaction of these products and the fuel, and as a result the gas turbine exhaust gases contain more carbon monoxide, hydrogen, soot and hydrocarbons than would be the case if the cooling airflow could be made discontinuous and made to participate in th flow within the combustion chamber. Also, the thick stream of relatively cool gas along the walls results in the temperture distribution of the efflux gases being less uniform than is desirable. Further, in the design stage of a combustion chamber, the compressor delivery air is considered in three broad divisions, burning air, cooling air and dilution air. With conventional cooling, the cooling air cannot be regarded as burning air or dilution air.In combustion chambers where the temperature rise is very high for performance reasons, so much air is required for burning that there may be insufficient air for colling and/or dilution. Thus, there is a need for the cooling air to not only cool but also to be used in the burning and dilution phases.

The present invention seeks to provide a combustion chamber arrangement which can be adequately cooled and in which the cooling air can be used in the burning and dilution phases.

According to the present invention there is provided gas turbine engine combustion equipment comprising a combustion chamber having fuel injection means at the upstream end, air inlet means for the inlet of compressed air and an open downstream end for the outlet of combustion products, the air inlet means including combustin and dilution air inlet means and cooling air inlet means located at axially spaced intervals along the combustion chamber, at least some of the cooling air inlet means being arranged to direct cooling air along the inner surface of the combustion chamber wall in both the upstream and downstream directions, the flow pattern generated by the combustion and dilution air being such as to entrain a substantial proportion of the cooling air.

The present invention will now be more particularly described with reference to the accompanying drawings in which: Figure 1 is a diagrammatic illustration of a bypass type engine incorporating one form of combustion equipment according to the present invention, Figure la illustrates in more detail the combustion equipment of Fig. 1, Figure 2 is a detailed view of the upstream end of the combustion chamber shown in Fig.

2, Figure 3 is a view to a larger scale of a cooling ring of the combustion chamber of Fig. 2, Figure 4 is a section on line 4-4 in Fig. 3 and, Figure 5 is a view on arrow 'A' in Fig. 4.

Referring to the drawings, a gas turbine engine comprises low pressure and high pressure compressors 1 2 and 1 4 respectively, driven by low pressure and high pressure turbines 1 6 and 1 8 respectively. A proportion of the compressor delivery air from the compressor 1 2 flows into a bypass duct and subsequently mixes with the exhaust gases from the turbine 1 6 in an exhaust duct 22.

The remaining compressor delivery air passes to the compressor 14, the delivery from which flows into combustion equipment 24 which also receives a supply of fuel (not shown).

The combustion equipment 24 is of the tubo-annular type in which a number of equispaced cylindrical flame tubes 26 are located in an annular casing 28.

Each flame tube, as shown in Fig. 1 a, comprises a head portion 30 in which is located a primary air inlet and fuel injector 32 of the type disclosed and claimed in our UK patent no. 1427146, and a number of cylindrical wall sections 34, 36 and 38 joined together by cooling rings 40, 42 and 44, the wall section 34 having a cooling ring 35 with air inlets 35a.

The flame tubes are joined together by interconnectors 46 and the wall portions 36 and 38 have inlets 48 and 50 for the inlet of secondary and dilution air respectively into the fame tube. Further inlets 52 and 54 of smaller diameter can be located in the planes containing the inlets 48 and 50 or they can be displaced either upstream or downstream of these planes to provide local cooling if required.

Each of the cooling rings 40 and 42 has at least one set of upstream directed cooling air inlets 40a and 42a and one set of downstream directed cooling air inlets 40b and 42b. Additionally, as shown in Figs. 2 to 5, themost upstream cooling ring 40 can have extra cooling air inlets 40c provided to prevent overheating of that part of the ring which extends furthest into the flame tube. This feature can also be applied to all the cooling rings if required. In order to encourage the cooling air to flow in both the upstream and downstream directions from at least the cooling rings 40, 42, a scoop ring 56 which is part of the wall section 34 is provided around the flame tube in the region of the cooling ring 40 to define an annular space 60 into which cooling air flows.

As shown in detail in Fig. 3, each of the air inlets 40a, b and C can be in the form of three rows of holes, the rows being staggered with respect to each other to provide adequate material between adjacent holes whilst main- - taining a close spacing.

The most downstream cooling ring 44 may only require an upstream directed set of cooling air inlets 44a, although cooling air inlets corresponding to inlets 40 b and 40 c can be provided if required. The general nature of the cooling rings 40, 42 and 44 and particularly that of the cooling ring 40 are disclosed in our co-pending published UK patent applicatin no. 2021204A.

In operation, fuel with air from the compressor 1 4 is injected using a co-axial dual orifice fuel nozzle (not shown) into a duct 32a where it is further atomised and partially vapourised on the hot surfaces of the duct 32a and a deflecting member 32b. The air and fuel mixture issuing from the injector creates a vortex 60 as described in UK patent no.

1427146 in zone A of the flame tube, and the upstream cooling air from the inlets 40a of the cooling ring 40 merge with this vortex and becomes combustion air. The next downstream zone, zone B is cooled by downstream cooling air from inlets 40b of the cooling ring 40. Cooling air from the Zone B is terminated by the ring of plunged holes 48 supplying combustion air, and a vortex 62 is formed which rotates in the opposite direction to the vortex 60, and it is in zone B that much of the reaction between the fuel and air takes place.

The cooling air from the inlets 40b and 35a merge with the flow in the vortex 62 and takes part in this reaction.

The wall of Zone B is cooled by air moving up stream from the inlets 42a of the cooling ring 42 and another vortex 64 is formed by part of the air flow through inlets 48 and 50, this vortex rotating in the same direction as the vortex 60. Further burning takes place in Zone C and since the cooling air from the inlets 42a merges with the vortex 64, this air can also take part in the combustion. It will be noted that no combustion products can be carrier downstream on the walls of Zone C.

Zone D is a continuation of Zone C and a further vortex 66 is generated by downstream cooling air from the inlets 42b for cooling the walls of Zone D and dilution air from the inlets 50.

The walls of Zone E are cooled by upstream air from the inlets 44a of cooling ring 40. In this zone and in regions downstream of this, burning is substantially completed and mixing of the hot combustion products with dilution air takes place.

At the position along the flame tube wall where the cooling air flows in opposite directions meet, there may be a tendency to overheat due to high turbulence at these positions.

Any such overheating can be dealt with by the provision of cooling air through the inlets 52 and 54 which will also have the effect of stabilising the position at which the cooling flows leave the flame tube walls and merge with the gaseous flow pattern within the flame tube.

It will be noted that by directing cooling air from a number of axially spaces locations along the flame tube in both upstream and downstream directions and by generatin a flow pattern within the flame tube which encourages these streams of cooling air to be detached from the flame tube walls, the cooling air can take part in the combustion and dilution processes. Such an arrangement will eliminate or reduce to an acceptable level, the tendency of the cooling air layer to grow in thickness as it moves downstream, and for the cooling air layer to carry unburnt fuel and combustion products. Further the cooling air will largely cease to be parasitic.

The invention has been described using a particular form of fuel injector but other forms can be used in conjunction with inlets for further combustion and dilution air which generate flow patterns of the type likely to entrain the cooling air flows on the walls of the flame tube.

Not all of the cooling rings need have upstream and downstream cooling air flows, particularly the most downstream cooling ring which may only be required to provide an upstream cooling air flow.

The invention has been described with. reference to a tubo-annular combustion system, but it can also be applied to systems of the multiple chamber or annular types.

Claims

CLAIMS "

1. Gas turbine combustion equipment comprising a combustion chamber having fuel injection means at the upstream end, air inlet means for the inlet of compressed air and an open downstream end for the outlet of combustion products, the air inlet means including combustion and dilution air inlet means and cooling air inlet means located at axially spaced intervals along the combustion chamber, at least some of the cooling air inlet means being aranged to direct cooling air along the inner surface of the combustion chamber wall in both the upstream and down stream directions, the flow pattern generated by the combustion and dilution air being such as to entrain a substantial proportion of the cooling air.

2. Combustion equipment as claimed in claim 1 in which some of the cooling air inlet means comprise a ring having a set of cooling air inlets extending in a generally upstream direction and a set of cooling air inlets facing in a generally downstrean direction.

3. Combustion equipment as claimed in claim 2 in which each set of cooling air inlets comprise at least two rows of inlet holes, the adjacent rows in each set being staggered with respect to each other.

4. Combustion equipment as claimed in any one of the preceding claims in which each cooling air inlet means includes a portion extending into the flame tube, the said portion of at least one of said cooling air inlet means including a set of inlets for the throughflow of cooling air.

5. Combustion equipment as claimed in any one of the preceding claims in which in use the flame tube, includes at least two regions of re-circulating flow, the cooling air inlet means being positioned so that cooling air becomes entrained in the re-circulating flow in the flame tube.

6. Combustion equipment as claimed in claim 5 in which the flame tube includes primary air inlet means at the upstream end of the flame tube, combustion and dilution air inlet means downstream of the primary air inlet means, and dilution air inlet means downstream of the combustion and dilution air inlet means, a first cooling air inlet means being located between the primary air inlet means and the combustion and dilution air inlet means, a second cooling air inlet means being located between the combustion and dilution air inlet means, and the dilution air inlet means and third cooling air inlet means downstream of the dilution air inlet means.

7. Combustion equipment as claimed in claim 6 in which the dilution air inlet means and the dilution air inlet means each comprise a ring of plunged holes in the flame tube, the rings being axially spaced apart.

8. Combustion equipment as claimed in claim 7 in which one or both rings of plunged holes, includes further holes of smaller diameter located between adjacent plunged holes.

9. Combustion equipment as claimed in any one of the preceding claims in which the fuel injection means is of the type described and claimed in our UK patent no. 1427146 or of the type described and claimed in our pending UK patent application no. 4558/77.

10. Combustion equipment as claimed in any one of the preceding claims in which at least some of the cooling air inlet means are of the type described in our pending UK patent application no. 2021204A.

11. Gas turbine engine combustion equipment constructed and arranged for use and operation substantially as herein described, and with reference to the accompanying drawings.