GB2158879A

GB2158879A - Preventing surge in an axial flow compressor

Info

Publication number: GB2158879A
Application number: GB08412865A
Authority: GB
Inventors: Rupert Morton Lucas
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1984-05-19
Filing date: 1984-05-19
Publication date: 1985-11-20
Also published as: IT1185000B; FR2564533A1; IT8520756A0; DE3517486A1; GB2158879B; JPS60256597A

Abstract

In axial flow compressors of gas turbine engines it is well known to use either an air bleed valve 34 or a compressor casing treatment to improve the performance of the compressor. Air bleed valves act to keep the working line of the compressor close to the surge line. The casing treatment, which usually consists of one or more grooves 28 in the internal cylindrical surfaces adjacent the tips of a stage of upstream blades 22, acts to raise the compression ratio at which the rotor stalls. This invention combines the use of the bleed valve 34 with the grooves 28 to allow the bleed valve setting point to be reduced and ensure the bleed valve is closed at low power settings. <IMAGE>

Description

SPECIFICATION Improvements relating to surge in an axial flow compressor This invention relates to gas turbine engines and more particularly to axial flow compressors for such engines. The Rolls-Royce Gem 4 is an example of a gas turbine engine to which the invention is applicable.

The value of airflow and pressure ratio at which a surge occurs is termed the 'surge point'. This point is a characteristic of each compressor speed, and a line which joins all the surge points, called the surge line (Figure 1), defines the minimum stable airflow which can be obtained at any rotational speed. A compressor is designed to have a good safety margin (Region A) between the airflow and the pressure ratio at which it will normally be opereted (the working line) and the airflow and compression ratio at which a surge will occur.

It is necessary to use a system of airflow control to ensure the efficient operation of an engine over a wide speed range and to maintain the safety margin referred to above. The system usually consists of bleeding compressor air from an intermediate stage either to atmosphere or, in the case of a bypass engine, into the bypass duct. The amount of air bled from the compressor may be controlled by mechanically, electrically or hydraulically operated valves which are open at low compression ratios and closed when the engine is operating at higher conditions. When the valves are open the airflow across the high pressure section is decreased, whilst the airflow across the earlier stages is increased.

This prevents the front stages from stalling because of a low axial velocity of air.

Bleeding compressor air to avoid stalling is wasteful of air and reduces the compressor's overall efficiency. An alternative solution may be to replace the bleed valve by introducing a form of compressor casing treatment, which consist of one or more recesses formed on the internal cylindrical surfaces adjacent the tips of a stage of rotors. Typical treatments are disclosed in U.S. Patent No. 4,239,452, U.S.

Patent No. 4,238,170, NASA Technical Notes TN D-6538, The American Society of Mechanical Engineers publications 75-GT-95 and 75-GT-13.

Such treatments have the effect of increasing the pressure ratio for any given mass flow at which surge actually occurs. To obtain maximum surge margin improvement from such treatment, complicated slot geometries are required, these can be costly, complex to machine and introduce as a modification and may have a power loss of up to 2%. A worthwhile improvement can still be obtained, however, by providing a simplified form of casing treatment as shown in Figure 3. Such treatment is much easier to machine and hence appreciably less expensive, but lacks the full benefit of the more costly option.

The Gem 4 engine is an uprated version of the Gem 2 having a mass flow increase of 1 5% over the latter, and is used in helicopters such as the Westland Aircraft Lynx and its derivatives.

Due to the increased mass flow rate, air is bled from the compressor by means of a Plessey fluidic controlled bleed valve. This valve is designed to operate at a predetermined value of the engine torque (Tq) in order that the engine may remain surge-free at low power settings.

It has been found that helicopters, operating at reduced all up weights, experience bleed valve operation during take-off. In twin engine helicopter installations the mismatch between engine power levels, when one engine's valve closes before the other engine's valve, casts doubt on whether the pilot is experiencing an engine or control system failure, or bleed valve movement. This in turn casts doubt on whether or not the take-off should be abandoned.

The above problem produces a requirement to reduce the bleed valve setting below the current predetermined value of Tq so that both valves are closed and the engines are matched on torque at take-off.

An object of this present invention is to provide a means to allow the bleed valve setting point to be reduced to a lower predetermined value of the engine torque to ensure the bleed valve is closed at take-off and thus avoid any mismatch between engine power settings. It consists basically of adding a simplified form of casing treatment to supplement the bleed valve operation.

The effect of such a casing treatment is to raise the pressure ratio at which surge occurs particularly at low mass flow values, this in turn allows the bleed valve setting point to be reduced without the compressor operating under surge conditions.

The invention will now be described in more detail, by way of example only, with reference to the accompanying drawings, in which: Figure 1 is a graph of pressure ratio against mass flow illustrating a typical safety margin between the surge line and the working line of the compressor.

Figure 2 shows a side elevation of a gas turbine engine having a broken away compressor casing portion disclosing a diagrammatic embodiment of the present invention.

Figure 3 is an enlarged view of the Ist stage rotor blade tip region of the compressor illustrated diagrammatically in Figure 3.

Figure 4 is a graph of engine pressure ratio (PR) against mass flow (MF) for a Gem 2 engine having no bleed valve or compressor casing treatment. The surge line (A), the work ing line (B) and the Ist stage rotor stall characteristics (C) can clearly be seen.

Figure 5 is a graph of engine pressure ratio against mass flow for a Gem 4 engine having both a bleed valve and a simplified form of compressor casing treatment. The effect of the casing treatment on the surge line (E) is illustrated by line (H), whilst-the effect of the closing and opening of the bleed valve on the working line (F) is illustrated at points (F1,F2) and (F3,F4).

Referring now to Figure 2 of the drawings, a gas turbine engine 10 comprises, in flow series, a low pressure compressor 12, a high pressure compressor 14, combustion equipment 16, a high pressure turbine 18, a low pressure turbine 20, the engine terminating in an exhaust nozzle 22. The low pressure compressor 12 and low pressure turbine 20, and high pressure compressor 1 4 and high pressure turbine 18 are each rotatably mounted upon a coaxially arranged shaft assembly not shown in the drawing. A diagrammatic view of an embodiment of the present invention is shown within the broken portion of the low pressure compressor casing at 1 3.

Figure 3 of the drawings shows a crnss- sectional view in greater detail of that shown diagrammatically in Figure 1. The drawing comprises a low pressure compressor blade 22 integraily mounted to a low pressure compressor shaft 24 and a portion of low pressure compressor casing 26 arranged radially outward of the low pressure compressor blade 22. Five grooves formed in the compressor casing 26, shown at 28, form the casing treatment. The grooves may for example be each Imm deep and Imm wide, arranged to be substantially perpendicular to the engine axis and equi-spaced along the low pressure compressor casing 26 above the low pressure compressor blade 22. Also illustrated is a ducting tube 30 provided in the compressor casing 26 radially outward of a second downstream compressor blade 32.A bleed valve shown schematically at 34 adjacent the tips of the second downstream compressor blade 32 is used to control the flow of bleed air from the compressor. Actuation means not shown but well understood per se are provided to open and close the bleed valve 34.

For satisfactory operation of a compressor stage such as that shown at 22 it is well known that it, and also the adjacent stages of blades must be carefully matched as each stage possesses its own individual airflow characteristics. It is, therefore, extremely difficult to design a compressor to operate satisfactorily over a wide range of operating conditions such as an aircraft engine encounters.

Outside the design conditions the gas flow around the blade tends to degenerate into a violent turbulence and the smooth pattern of flow through the stage or stages is destroyed.

The gas flow through the compressor usually deteriorates and the stalled gas becomes a rapidly rotating annulus of pressurised gas above the tips of one compressor blade stage or group of stages. If a complete breakdown of flow occurs through all the stages of the compressor such that all the stages of blades become 'stalled', the compressor will 'surge'.

The transition from a 'stall' to a 'surge' can be so rapid as to the unnoticed or on the other hand may be so weak as to produce only a slight vibration or poor acceleration or deceleration characteristics. A more severe compressor stall is indicated by a rise in turbine gas temperature, and vibration or coughing of the compressor. A surge is evident by a bang of varying severity from the engine compressor and a rise in turbine gas temperature.

As stated above, both bleed valves and simplified compressor casing treatments individually provide a degree of control in allowing operation of the compressor at high pressure ratios without experiencing surge problems. Combining the bleed valve with a simplified compressor casing treatment in a previously unthought of combination meets the particular requirement to allow reduction of the bleed valve setting point to the lower predetermined value of engine torque and thus ensure the valves are closed during takeoff.

Figure 4 is a graph of pressure ratio against mass flow for the compressor of a Gem 2 engine having neither a bleed valve or compressor casing treatment. The line A represents the surge line of the compressor, line B represents the 'working line' of the compressor whilst line C represents the first stage rotor stall characteristics. As detailed above, there is a requirement to keep the working line (B) as close to the surge line (A) as possible whilst still maintaining a safety margin (Region D). It can be seen from Figure 4 that as the mass flow through the compressor increases the surge line and the working line diverge with the consequence that the effectiveness of the compressor is reduced. This loss in effectiveness is greater still on the Gem 4 because of its 15% increase in mass flow over the Gem 2, this lead to the introduction of a bleed valve 34 in an attempt to keep the working line of the compressor (B) as close as possible to the surge line of the compressor (A).

Figure 5 is a graph of pressure ratio against mass flow of the first stage rotor of the compressor for a Gem 4 engine having both a bleed valve (34) and the simplified form of compressor casing treatment (28). The line E represents the surge line, line F represents the 'working line', line G represents the first stage rotor stall characteristics and line H illustrates the effect of the simplified compressor casing treatment on the surge line of the compressor.

From the working line (F) the characteristics of the bleed valve operation can clearly be seen.

As detailed above when the mass flow through the compressor increases the working line and the surge line diverge, the bleed valve (34) is designed to close between points F1 and F2 on the graph and introduce a 'step' in the working line in order to bring it closer to the surge line. As the mass flow decreases the bleed valve opens between points F3 and F4, in order to avoid operating the compressor under surge conditions. Due to the stepped opening and closing characteristics of the bleed valve it is presently impossible to reduce the bleed valve setting point to below engine torque (Tq) without encountering a surge problem.

The simplified compressor casing treatment raises the pressure ratio at which surge occurs for any given mass flow rate, as illustrated by line of the graph. This raising effect is sufficient to allow a reduction of the bleed valve setting point to the lower predetermined value of engine torque whilst still maintaining a safety margin between the surge line and the working line of the compressor. At this setting point bleed valve closure is ensured on both engines during take-off thus avoiding any mismatch between engine power levels.

A further effect of the casing treatment is to alter the stalling characteristics of the first stage rotor, as illustrated by line G.

The invention as described above has a further benefit relating to low pressure first stage compressor blade failure. Investigations have indicated that stress caused by blade flutter is generated due to inadequate bleed valve mass flow at or near the zero power auto rotative decent conditions. This stress is sufficient to cause cracking of the blades and could lead to blade failure causing very extensive secondary damage. One solution to this problem would be to increase the bleed valve flow. However, this would also have the effect of increasing the amount of torque mismatch between engines, as described above. The exploitation of the present invention will raise the surge line at zero bleed conditions sufficiently to control the blade flutter stress to below that required for effectively infinite life.

Claims

1. An axial flow compressor, comprising, a casing in which is mounted a rotor carrying at least one row of generally radially extending blades, characterised in that the casing is provided with both a number of recesses disposed within its internal cylindrical surface adjacent the tips of at least one blade row and an air bleed valve disposed to bleed air from the compressor downstream of the one or more recesses.

2. An axial flow compressor according to Claim 1 in which the recesses consist of five circumferentially extending grooves.

3. An axial flow compressor according to Claim 2 in which each groove is 1 mm deep, 1 mm wide and is arranged to be substantially perpendicular to, and equi-spaced along, the compressor longitudinal axis.

4. An axial flow compressor according to Claim 1 or Claim 2 in which the air bleed valve is arranged to operate at rotor speeds which may result in the compressor surging due to adverse airflow characteristics and closed at all other engine speeds.

5. An axial flow compressor according to any of Claims 1 to 4 in which the effect of the recesses is to raise the airflow and compression ratio at which the compressor will surge over a range of compressor speeds.

6. An axial flow compressor according to any of Claims 1 to 5 in which the effect of opening the bleed valve is to increase the safety margin between the airflow and compression ratio at which the compressor works and the airflow and compression ratio at which the compressor surges.

7. An axial flow compressor according to Claims 1 to 5 in which the recesses act in conjunction with the bleed valve to allow a reduction in the airflow and compression ratio at which the bleed valve operates.

8. An axial flow compressor as claimed in any of the preceding claims substantially as herein described by way of example only and with reference to Figures 2, 3 and 5.