GB2208894A - Gas turbine engine compressor stall recovery - Google Patents

Abstract

In an arrangement for recovering a gas turbine engine from a compressor stall condition, the effective outlet area of the compressor is temporarily reduced. This reduces the overall working line of the engine sufficiently to promote recovery from the stall condition whilst avoiding the problems of a flame-out which is an inherent hazard in the conventional cure of fuel-dipping. Several proposals for achieving the area reduction of the outlet 8 of the compressor 2 are described involving, an axially translatable plug 10, a variable area exit nozzle (Fig. 2) and rotatable exit vanes (Fig. 3). <IMAGE>

Description

IMPROVEMENTS RELATING TO A GAS TURBINE ENGINE The invention relates to a gas turbine engine and, in particular, relates to means for assisting recovery of an engine from a compressor stall condition.

In a gas turbine engine air entering the engine is compressed before combustion takes place. The expanding gas stream then passes through the turbine which is usually coupled directly, via a turbine shaft, to drive the compressor. The compressor rotor is thus turned at high speed so that air is continuously induced into the engine through the compressor inlet and swept rearwards usually, as for example in an axial flow compressor, through several compressor stages.

The pressure of the air increases progressively as it diffuses rearwards through the rotor and interdigitated stator stages. The stator blades also serve to correct the deflection imparted to the airflow by the rotor blades. The last stage of stator blades, otherwise referred to as compressor outlet guide vanes act to "straighten" the airflow so that it enters the combustion chamber, or chambers, substantially free of swirl and at a reasonably uniform velocity.

Each compressor stage produces a relatively small pressure increase so as to limit losses due to the airflow breaking away from the blades and thereby inducing subsequently a blade stall condition.

The compressor stages may be forced to deliver a higher pressure rise than design, either uniformly or locally, by such typical causes as increased fuel flow rates during acceleration or zones of intake pressure depression associated with aircraft operation at a high angle of attack. If the deviation from design conditions is sufficient, the increased compressor blade incidence which is induced by increased backpressure, leads to separation of airflow from the blade surfaces, otherwise known as stalling.

If the extent of stalling is sufficient, continued operation of the compressor in a high pressure rise mode becomes unstable. One or the other of two different types of behaviour commonly follow the onset of such an instability; First, the pressure rise almost collapses, permitting noisy flow reversal (a "surge") and relieving the pressure in the core of the engine with consequent compressor recovery. After recovery, backpressure builds up again to return the unit to its initial instability point and therefore often establishing a violent cyclic disturbance to engine performance.... commonly known as "cyclic surge".

Second, the pressure rise falls dramatically but does not cause a general flow reversal (and therefore makes little noise). It is found that the engine may then run in a state in which the compression system is passing flow only in a segment of its annulus, whilst the complementary segment contains churning stalled flow of near-zero throughflow velocity. The segments always rotate and, in this state, generally affect the whole annulus height.... a condition known as "full span rotating stall" in the compressor, whilst the resulting deteriorated engine state is often known as "stall", "locked stall" or "hang-up".

The consequences of engine stall are frequently an almost complete loss of thrust and a severe rise in flame temperature and under some circumstances may lead to engine overtemperature damage even though the control systems are responding in their intended fashion. It is found that cyclic surge events can usually be interrupted and corrected by causing a rapid decrease in fuel flow to the engine, thus preventing the demand for backpressure returning to the level which precipitated the event. Engine stalls prove to be much more difficult to clear. It is typically found that even a total cancellation of fuel supply is insufficient to cause prompt recovery...

that is, the working line level required for recovery is lower than the engine unlit working line. In such cases, the engine can often only be restored to normal operation by shutdown and deceleration to extremely low speed... in flight windmilling has been known to keep the engine unlit speed above that needed for recovery. There are circumstances therefore, where engine stall can result in aircraft loss if that aircraft has only a single engine.

It is, therefore, an object of the present invention to provide means capable of recovering an engine from a full-span stall condition without engine shut-down.

According to the present invention there is provided a gas turbine engine comprising a compressor for compressing air entering the engine prior to combustion, the compressor having an outlet and tetricting means for temporarily reducing the effective area of the compressor outlet, said restricting means normally occupying a first position in which the compressor outlet is unrestricted and having a second position in which the compressor outlet is partially retricted by a substantial reduction in the orifice area of the compressor outlet and actuating means for deploying the restricting means from said first position to said second position for a brief period of time to alleviate a compressor stall condition.

Preferably the compressor outlet is defined by a duct and the means for temporarily reducing the effective outlet area comprises means mounted within the duct, the means being mounted for movement from a first position which defines a first effective outlet area to a second position which defines the reduced outlet area.

In one embodiment of the invention the means for reducing the compressor outlet area comprises axially translatable plug means movable into an outlet duct of the compressor.

In another embodiment the means of reducing the compressor outlet area comprises variable area nozzle means.

In a further embodiment a plurality of rotatable guide vanes are monted in a compressor exit duct, said vanes being rotatable in unison, each about its own longitudinal axis to reduce the compressor outlet area.

The invention will now be described by way of example with reference to the accompanying drawings, in which: Fig 1 is a diagrammatic illustration of an axially translatable compressor outlet plug, Fig 2a is a diagrammatic illustration of a variable area compressor outlet nozzle, Figs 2b and 2c show details of alternative nozzle arrangements, Fig 3 is a diagrammatic illustration of a compressor outlet duct provided with vanes variable by rotation of the whole airfoil or of a flap forming the rear section of the airfoil, and Fig 4 shows a compressor performance diagram demonstrating the improvement conferred by the invention.

The three examples of embodiments of the invention illustrated in Fig 1, 2 and 3 concern multi-stage, axial flow compressors of the same type. In the drawings like parts are given like references.

Referring first to Fig 1, the compressor inlet is shown to the left of the drawing at 2, the rotor and stator assemblies are indicated generally at 4, and the combustion chamber assembly at 6 towards the right of the drawing. Disposed between the aero-mechanical section 4 of the compressor and the combustion chamber region 6 lies an annular duct 8 comprising the compressor outlet. Generally duct 8 is designed to maintain a steady airflow velocity and pressure in the compressor outlet.

An annular plug 10 is mounted in the duct 8 concentric with the core of the engine for movement in an axial direction towards and away from the rear of the compressor section 4. The plug 10 co-operates with the wall of the duct 8 to define the effective compressor outlet area. The plug is translatable . axially relative to a longitudinal axis of the engine whereby the plug co-operates with the wall of the duct to define a progressively reducing outlet area as the plug is moved from its first position to its second position, shown by the dashed line in Fig 1. As illustrated the plug 10 has a forwardly tapered cross-section. The duct 8 is co-axial with the core of the engine and thereby defines an annular outlet from the compressor. The plug 10 is also annular.

Means is provided for driving the plug and for controlling its location within the duct 8. The cross section of the plug is designed to meet two basic criteria; first, that it shall not unduly affect the characteristics of airflow in duct 8 when in its normal position, ie when the compressor is running within its design limits in an unstalled condition, and second, that when deployed, ie moved towards the compressor, it shall bring about a significant reduction in the area of the compressor outlet.

Operating means similar to that used for controlling movement of a known exhaust nozzle plug may be arranged to move the translating plug 10 from its normal to its deployed position for relatively short periods of time only, as and when required to cure a stall condition. The plug 10 is to be deployed for short periods only, typically of the order -of 0.1 sec. However, shorter or longer periods of restriction than this may be required to cure compressor stall conditions in engines of different sizes depending on engine size, mass flow and pressure ratio etc.

The illustrated example shows a plug 10 of generally parallelogram cross-section although other shapes of plug may be employed. It is the shape of the leading edge or surface, that is the side facing the compressor, which is the more important as this remains exposed when the restricting plug is withdrawn to its normal position. Its exposed face should remain aerodynamically efficient so that corners and sharp edges which might disrupt airflow are to be avoided.

The second illustrated embodiment of Fig 2 comprises a contractable or variable area annular nozzle 12 located adjacent the rear annular orifice of the compressor section. The nozzle 12 comprises inner and outer concentric rings 14,16 of interlocking flaps. The outer nozzle ring 14 comprises a plurality of circumferentially spaced flaps hinged to the rear of the compressor casing around its outer periphery. The inner nozzle ring 16 also comprises a plurality of flaps spaced apart around the inner circumference of the compressor annular outlet and hinged there to the casing.

The nozzle flaps are preferably shrouded by the duct 8, or otherwise sealingly attached to the rear of the compressor casing in order to avoid pressure losses in the compressor output due to gas leakage. The two rings of flaps are actuated in a similar manner to that used to operate turbine exhaust nozzle flaps. For example the flaps may be operated by.screw jacks to the closed or restricted position and by gas pressure to the open or normal position. The flaps may also be pneumatically operated.

Each of Figs 2b and 2c show an alternative variable area nozzle arrangement utilising only one ring of pivoted flaps which co-operate with a fixed wall of the duct 8 to define a variable outlet area. Fig 2b shows a detail of an arrangement using the inner ring 16 of movable flaps, and Fig 2c shows a corresponding illustration of an arrangement using the outer ring 14 of movable flaps.

The third arrangement shown in Fig 3 basically comprises a plurality of variable exit guide vanes 18 equi-angularly spaced around the annular duct 8 close to the rear of the compressor section.

During unstalled operation, the outlet guide vanes (18) are positioned as shown by the solid line detail in the drawing. In order to give a much reduced effective outlet area to aid recovery from stall, the outlet swirl can be increased by rotation of the whole of the airfoil or of a flap forming the rear section of the airfoil (both configurations'are indicated in the dashed lines in Fig 3). The amount of rotation required would usually lie between 300 and 600 and would normally be in a direction to reduce the total amount of turning required of the vane row. The vanes are of known airfoil section.

The vanes 18 are mounted in the aft end of the compressor casing or in the duct 8 each for rotation about its longitudinal axis. Each vane is attached at its radially outer end to a lever 20 which is linked to a circumferentially rotatable unison ring 22 mounted in a plane parallel to the plane of the vanes and perpendicular to the axis of the engine. The same type of arrangement is known for controlling variable guide vanes in other parts of an engine, eg movable inlet guide vanes. The unison ring is driven through a pinion or recirculating screw arrangement by a motor.

Downstream of the variable vanes further fixed vanes 24 may be optionally provided to act as flow straighteners for air exiting the duct 8 into the combustion chamber region 6. If fitted, such straighteners (24) must be at a significant distance from the outlet guide vanes (18) in order to permit the flow to approach static pressure uniformity at the outlet guide vane exit, a representative distance would be about twice the radial height of the annular duct (8).

In operation, all three described arrangements function essentially in the same way to reduce the static pressure level at the compressor exit orifice by reducing the actual (as in Figs 1 and Fig 2) or effective (as in Fig 3) outlet area.

The most common stall condition found is a full-span rotating stall consisting of a single rotating sector of near-zero axial flow velocity, although near-normal flow continues around the remainder of the compressor. Pressure in the stalled sector or cell produces a significant pressure rise virtually independent of sector size making the exit total pressure substantially equal to the static pressure at that point. Where there is negligible airflow, as in a stalled sector, the static pressure at the outlet is unaffected by a change in the outlet area. But, in the case of a uniformly flowing region the limiting value of static pressure at the compressor exit falls in proportion to the exit area. Thus, this lowering of static pressure by temporarily reducing the exit area is effective in promoting stall recovery.

Referring now to Fig 4, normal unstalled operation of the full 360 degrees of the compressor is represented in terms of the ratio of outlet total pressure/inlet total pressure versus total flow by curve ABCD. Deep stall 360 degrees operation ie at zero total flow is represented by point K at which the outlet total and static pressures are equal. The normal characteristic re-evaluated in terms of outlet static pressure/inlet total pressure is given by curve EFG.

If the compressor is forced to operate with a throttle size such as represented by the demand line OPQ, the machine has to operate partly in stall at point K and partly unstalled at the equal outlet static pressure point F, this will give a mean operating point P. If P were halfway along KF, the compressor would have a 180 degree stall cell. The point S is the stall dropout point drawn on the assumption that the critical minimum size of a stall cell is 90 degrees. The equivalent operating point expressed in average total pressure terms (assuming the useful total pressure at outlet would be the area mean) is shown by point U lying on the line KC where C is the local total pressure rise value corresponding to operation at F.

Reduction of the compressor outlet area, keeping the same compressor internal geometry, would leave ABCD unaltered but would drop the static pressure rise to curve HIJ. In stalled operation the unstalled zones would operate at point I and the stall dropout point (90 degree cell) would move to R, and the equivalent total pressure point to B.

The average dropout point will then become T, assuming that the critical minimum stall cell size remains at 900. The increased pressure rise and reduced flow of T relative to U defines the rise (improvement) in working required for dropout.

The increased diffusion required as a result of decreased outlet area will usually result in increased downstream total pressure losses detracting from this improvement, but normally these losses can be kept relatively small.

Claims (16)

1. A gas turbine engine comprising a compressor for compressing air entering the engine prior to combustion, the compressor having an outlet, and restricting means for temporarily reducing the effective area of the compressor outlet said restricting means normally occupying a first position in 'which the compressor outlet is unrestricted and having a second position in which the compressor outlet is partially restricted by a substantial reduction in the orifice area of the compressor outlet and actuating means for deploying the restricting means from said first position to said second position for a brief period of time to alleviate a compressor stall condition.

2. A gas turbine engine as claimed in claim 1 wherein the compressor outlet is defined by a duct and the means for temporarily reducing the effective outlet area comprises means mounted within the duct, the said means being mounted for movement from a first position which defines a first effective outlet area to a second position which defines a second reduced outlet area.

3. A gas turbine engine as claimed in claim 2 wherein the means for reducing the outlet area comprises plug means which co-operates with a wall of the duct to define the effective compressor outlet area, and the plug means is movable relative to the duct between the said first and second positions.

4. A gas turbine engine as claimed in claim 3 wherein the plug means is translatable axially relative to a longitudinal axis of the engine in order to define co-operatively with the duct a progressively reducable outlet area.

5. A gas turbine engine as claimed in claim 4 wherein the duct defines an annular outlet area and the plug means is also annular.

6. A gas turbine engine as claimed in claim 5 wherein the plug means has an axially tapered cross-section.

7. A gas turbine engine as claimed in claim 2 wherein the means for reducing the effective outlet area of the compressor comprises a variable area nozzle.

8. A gas turbine engine as claimed in claim 7 wherein the variable area nozzle comprises a plurality of flaps spaced circumferentially around the compressor outlet, the flaps -being pivotally mounted and movable from a first position which defines a first effective outlet area to a second position which defines a second reduced outlet area.

9. A gas turbine engine as claimed in claim 8 wherein the flaps are arranged in two concentric rings whereby to define between them an annular compressor outlet of variable effective area.

10. A gas turbine engine as -claimed in claim 8 wherein the movable flaps are arranged to co-operate with a fixed wall of the duct to define the variable area compressor outlet.

11. A gas turbine engine as claimed in claim 2 wherein the means for reducing the effective outlet area of the compressor comprises movable exit guide vanes mounted in the duct defining the compressor outlet.

12. A gas turbine engine as claimed in claim 11 wherein the guide vanes are spaced radially at equal intervals around the outlet duct, each of the said vanes has a substantially aerofoil shaped transverse section, and is mounted for rotation about a longitudinal axis whereby each of the said vanes is movable from a first position in which they define a first effective outlet area to a second position in which they define a second reduced outlet area.

13. A method of alleviating a compressor stall condition in a gas turbine engine comprising the step of partially restricting the orifice area of the compressor outlet for a brief period of time.

14. A method according to claim 13 wherein means capable of substantially reducing the compressor outlet area is deployed to a restricting position for a brief period of time in response to the detection of incipient stall conditions.

15. A method of alleviating a compressor stall condition substantially as described with reference to the accompanying drawings.

16. A gas turbine engine substantially as described with reference to the accompanying drawings.