GB2212607A - Combustion chamber for a gas turbine engine - Google Patents

Abstract

A gas turbine engine combustion chamber has one or more wall-cooling devices 10 each comprising a transversely extending combustion casing wall lip 11, defining a sheltered and constrained gas space 12 on the frame side of the casing, and cooling air inlets 13 on the lee-side 12 of the lip aligned to cause an adherent film of cooling air to form by impingement on the casing. The air flow from the inlets has an upstream component which passes into the space 12 and is redirected in a downstream direction. <IMAGE>

Description

WALL COOLED GAS TURBINE ENGINE COMBUSTION CHtHBER This invention relates to a gas turbine engine combustion chamber having improved arrangements for cooling its wall materiai.

It is conventional in gas turbine engines for the combustion chamber to comprise a double walled structure. Tne outer structure is termed the pressure casing. Within this there is a structure termed the liner or alternatively the flame casing. The liner serves to confine the combustion gases and defines within it the combustion zone. The pressure casing is an envelope into which the compressed air delivered by the compressor is directed. There is an air-space between the liner and the pressure casing and a portion of the compressor delivery air passes through this air-space. It is conventional to utilize the compressor delivery air available within this air-space as a coolant for cooling tne wall of the liner.

The basic structure described aDove is found in all three established varieties of gas turbine combustion chambers: the multiple chamber type; the tubo-annular type; and the annular type. The description and explanation given in this specification assumes a conventional configuration of the double walled type.

Cooling is required for the liner walls in order that the wall material can survive for a useful lifetime in the harsh environment of the combustion zone. The quest for ever improved thermodynamic efficiency and/or improved thrust to weight ratio tends to exacerbate the difficulty of cooling the liner wall adequately. Higher turbine entry temperature, which is desirable for improved thermodynamic efficiency, requires a minimal elf flux of spent wall cooling gases into the combustion gases. Higher compressor pressure ratios raise the temperature of the compressor delivery air and reduce its cooling capacity.

This invention is concerned with film cooling, which is a technique by which the liner wall is covered on its flame side by a film of compressor delivery air so that it is cooled by the air in the film and is also separated from contact with the hot gases thereby. This technique is well established in the art as are others such as forced convection cooling, impingement cooling, transpiration cooling, and pseudo-transpiration cooling. A general description of these techniques is given in a paper by A B Was sell and J K Bhangu entitled "The Development and Application of Improved Combustor Wall Cooling Techniques" which was presented in 1980. This paper is available under ASME reference 80-GT-6O.

Despite the best that the current state of art can offer by way of combustor liner wall-cooling, thermal degradation of the liner is a life limiting factor for present day military aero engines at least and the problem will be worse in projected engines of the future unless some breakthrough in wall cooling is achieved.

Prior art film cooling techniques have been less effective than might be desired in part because the cooling film rapidly loses its effectiveness in its travel down the combustion chamber through growth of a sluggish boundary layer. The presence of such a layer causes shear-induced turbulence within the cooling film and promotes entrainment of hot gases into the cooling film.

This invention provides a novel combustion chamber geometry aimed at the provision of a wall cooling film of increased effectiveness.

The invention is a gas turbine engine combustion chamber having at least one device therein for cooling the liner of the combustion chamber, said device comprising: a lip on the flame side of the liner which extends transversely of the longitudinal axis of the combustion chamber thereby defining a sheltered lee-side zone adjacent the downstream surface of the lip; and an array of air outlets disposed around the lip, each air outlet being set in or adjacent the downstream surface of the lip and spaced from the root thereof within the lee-side zone, and each air outlet being directed such that, in operation of the device, a jet of air is caused to impinge upon the liner within the lee-side zone at an angle of incidence such as to ensure there is an upstream component in the dispersal flow;; and wherein the overall configuration of the device is such that the downstream surface of the lip constrains the upstream component of the dispersal flow and redirects it in a downstream direction to combine with the downstream component of dispersal flow thereby to produce a film of air flowing downstream adjacent the flame side of the liner.

The cooling film produced by the array of air outlets, which includes the downstream component of the dispersal flow and the redirected upstream component of the dispersal flow is fundamentally vortical in nature. Each jet of cooling air gives rise to a pair of linked vortices, one each side of the jet in the transverse plane, which are swept downstream as spiral flows in the cooling film. This flow pattern is resistant to boundary layer growth because air adjacent the flame side surface of the liner is re-energised by the spiral flows within the film. Moreover, it is postulated that the rotational motion of the vortex pairs is such as to enhance the adherence of the cooling film to the liner surface.

The configuration of the cooling device in the claimed combustion chamber and the nature of the flow pattern within the cooling film created by this device distinguish the invention from prior art combustion chambers and are responsible for its beneficial performance. Prior art combustion chambers utilize either discharge of the cooling air directly in the downstream direction parallel to the flame side of the liner, or inwardly-directed jets of cooling air onto splash plates which divert the air downstream. Neither of these arrangements enhances the resilience and life of the cooling film in the way achieved by the invention since they do not deliver cooling air with a radially outward component and at an angle of incidence which induces some upstream flow.

In a preferred arrangement the lip of the cooling device is so configured as to appear in section as a promontory aligned in the general direction of the outlet end of the combustion chamber so that it overlies and thus confines the gases at the upstream end of the sheltered lee-side zone.

In one arrangement the lip of the cooling device sits within a transversely extending recess in the combustion liner so that the upstream surface of the lip does not protrude beyond the general line of the combustion liner and so that it does not extend into the hot gases within the combustion chamber.

It is convenient for the array of air outlets to comprise a single row of regularly spaced outlets disposed around the lip. There might, however, be advantage in an array of two or more parallel rows with staggered spacing. It is important for the best operation of the device that the outlet spacing in the array is sufficiently close that the downstream flows resulting from adjacent jets of cooling air expand to meet each other in the transverse direction before they leave the sheltered lee-side zone, in order to avoid the possibility of entraining hot gases from the combustion chamber.

Exemplary forms of the invention are described below with reference to the drawings, of which: Figure 1 is a partially sectioned isometric projection of a combustion chamber; Figure 2 is a partial longitudinal section of the combustion chamber illustrated in Figure 1, depicting the configuration of the cooling device; Figure 3 is part transverse section view showing the flowfield created downstream of the cooling device; Figure 4 is a sectional view, similar to Figure 2, but showing an alternative form of cooling device; Figures 5 and 6 are partial transverse sectional views depicting alternative forms of air outlets; and Figure 7 is a partial transverse sectional view depicting an alternative array of air outlets.

Figure 1 depicts a representative combustion chamber of a gas turbine engine of the multiple combustion chamber type. In an engine of this type there is an annular array of separate combustion chambers having individual liners and individual pressure casings. In Figure 1 the combustion chamber is given a general designation 1, the pressure casing is designated 2 and the liner designatea 3. Combustion chamber 1 has an inlet 4, which connects to the exit annulus of the compressor section, and an outlet 5 which connects to the inlet annulus of the turbine section. At the inlet 4 there is a snout 6 which serves to divide the incoming airflow into two parts. That part of the airflow which is captured within the snout 6 is introduced into the primary zone of the combustion chamber for combustion purposes.The remaining part of the incoming airflow is directed into an airspace 7 which exists between the liner 3 and the pressure casing 2. The liner 3 is sealed to the pressure casing 2 at the outlet end of the combustion chamber so all of the airflow entering airspace 7 is passed to the interior of the combustion chamber 1. A major portion of this airflow passes through port entries 8 to be used in stabilizing the combustion, diluting the hot combustion gases etc as in conventional combustion chambers. The remaining portion of the airflow is utilized in cooling the wall of the liner 3.

The liner 3 incorporates cooling devices 10. Each of these is in the form of a lip 11 on the flame side of the liner 3 which extends around the liner wall. The lip 11 is transverse to the longitudinal axis of the flow of gases within the combustion chamber. In consequence each lip defines a lee-side zone, designated 12, which is in the wake of the lip and sheltered from the main stream of hot combustion gases.

Each lip 11 is in the form of a fold made within the wall of the liner 3. These can be most easily incorporated in the construction of the liner 3 at places where adjacent sections of liner are joined, in place of the corrugated overlap joint found in some known combustion chambers. On the downstream surface of each lip 11 there is a single row array of cooling outlets 13. These are closely and regularly distributed along the lip 11 and are inwardly spaced from the root of the lip 11.

Each of the outlets 13 is a simple round hole in the downstream surface of the lip.

More detail of the form and operation of the cooling device 10 is discernible in Figure 2. The lip 11 is a simple, generally triangular promontory on the flame side 21 of liner 3. The lip 11 establishes the sheltered lee-side zone 12 which is bounded by the downstream surface 22 of the lip 11 and an adjacent section 23 of the liner wall. The gas pressure to the flame side of the liner 3 is less than the air pressure within airspace 7. In consequence air issues from the outlets 13 as jets which cross the lee-side zone to impinge upon section 23 of the liner wall. This jet flow is dispersed upon impingement, partly laterally, partly in the upstream direction, and partly in the downstream direction.The magnitude of the upstream component of this dispersal flow, relative to that of the downstream component, is dependent upon the angle at which the jet impinges on the surface at section 23 of the liner wall. It is essential that there is an upstream component to this dispersal flow for such a component ensures that the region at the root of the lip 11 is not filled with stagnant trapped gases but is continually purged by the cooling air, and also because the upstream component of dispersal flow, upon redirection by the downstream surface 22 of the lip, serves to reinforce the flow pattern which is desired in the cooling film.

Figure 3 illustrates the flowfield which is created downstream of each cooling device 10. The jet 24 which issues from each outlet 13 is dispersed upon impingement with section 23 of the liner wall with components of dispersal flow in all outward directions. The upstream component of dispersal flow is restricted by the downstream surface of the lip 11 and is redirected in the downstream direction. The lateral components of dispersal flow are blocked on reaching the dispersal flow of neighbouring jets and in consequence these, too, are swept downstream. The resultant flowfield seen downstream of the outlets 13 as depicted in Figure 3 is one in which there is a vortex pair established by each jet 24 and this is swept downstream as spiralling flows. These spiral flows expand laterally until constrained by adjacent flows so that there is created a saturated field of double vortices.

The air which flows downstream in this field thus establishes a film of cooling air on the flame side of the liner 3. As the flow within the film is of a recirculating spiral pattern the growth of a sluggish boundary layer is resisted. The rotational sense of each vortex pair centred upon a respective jet 24 is such that it is believed to press itself against the liner wall by reactive force.

An alternative form of cooling device 10 is depicted in Figure 4. In this device the lip 11 sits within a traversely extending recess 30 in the wall line of the liner 3. This arrangement is intended to ensure that the upstream surface of the lip designated 31 and the tip portion 32 of the lip do not project into the main stream of hot combustion gases in the combustion chamber 1. To be most effective in this regard neither surface 31 nor the tip 32 should project beyond the general line of the combustion chamber liner 3. This general line is designated 33 in this figure.

Figure 5 depicts an alternative form of outlet 13. In this construction the outlets have a major axis and a minor axis the former being significantly greater than the latter and being set substantially normal to the surface of the liner 3 when seen in transverse section as depicted.

Figure 6 depicts an alternative array of outlets 13 consisting of two staggered rows. This arrangement will permit closer circumferential spacing of outlets 13 if minimal spacing in a single row array is limited by physical strength factors.

In an arrangement such as this it may be necessary for the outlets in respective rows to be of different sizes in order to produce a transversely saturated film of uniform profile, since jets issuing from the downstream outlets will have less opportunity for lateral expansion.

It has already been stated that there should be an upstream component to the dispersal flow resulting from impingement of the jets 24 on the section 23 of the liner wall.

This criterion imposes a limitation upon the tolerable angle of incidence. If the angle of incidence is defined as the angle between the jet and the normal to the wall at the point of incidence, when considered in longitudinal section as depicted in Figure 2, then the limiting value as predicted by theoretical studies will be around 510. In a practical configuration this theoretically derived value is unlikely to be definitive. However the requirement that there is an upstream component of dispersal flow might pose some problems in achieving a satisfactory jet angle together with a satisfactory lip/liner wall configuration. In the simple arrangements described above, for example, the jet angle is dependent on the lip/liner wall configuration. This limitation can be circumvented by use of directional nozzles in place of plain hole outlets as shown in the drawings. The directional nozzles may be set in the downstream surface 22 of the lip and supplied as previously supplied or may alternatively be mounted upon a cooling air gantry close to the downstream surface 22. In the latter case a separate supply from the compressor delivery is necessary.

Claims (8)

1. A gas turbine engine combustion chamber having at least one device therein for cooling the liner of the combustion chamber, said device comprising: a lip on the flame side of the liner which extends transversely of the longitudinal axis of the combustion chamber thereby defining a sheltered lee-side zone adjacent the downstream surface of the lip; and an array of air outlets disposed around the lip, each air outlet being set in or adjacent the downstream surface of the lip ane spaced from the root thereof within the lee-side zone, and each air outlet being directed such that, in operation of the device, a jet of air is caused to impinge upon the liner within the lee-side zone at an angle of incidence such as to ensure there is an upstream component in the dispersal flow;; and wherein the overall configuration of the device is such that the downstream surface of the lip constrains the upstream component of the dispersal flow and redirects it in a downstream direction to combine with the downstream component of the dispersal flow thereby to produce a film of air flowing downstream adjacent the flame side of the liner.

2. A gas turbine engine combustion chamber as claimed in Claim 1 wherein the outlets are arranged in a single row around the circumference of the combustion chamber on the lee-side of the lip.

3. A gas turbine engine combustion chamber as claimed in Claim 1 wherein the outlets are arranged in two or more staggered rows around the circumference of the combustion chamber on the lee-side of the lip.

4. A gas turbine combustion chamber as claimed in Claim 3 wherein the outlets in respective staggered rows are unequal in size.

5. A gas turbine engine combustion chamber as claimed in any preceding claim wherein the lip is a fold in the combustion liner wall and the outlets are holes in the lee-side surface of the fold.

6. A gas turbine engine combustion chamber as claimed in any preceding claim wherein the lip is disposed within a transversely extending recess in the combustion liner.

7. A gas turbine engine combustion chamber as claimed in any one of Claims 1 to 4 wherein the outlets are directional nozzles in or adjacent the lee-side of the lip.

8. A gas turbine combustion chamber as claimed in Claim 1 substantially as hereinbefore described with reference to Figures 1 to 3 of the accompanying drawings.