GB2108586A - Gas turbine engine active clearance control - Google Patents

Abstract

Clearance control between the tips of turbine rotor blades and the surrounding casing is effected by a valve 54 which, according to engine conditions, mixes relatively lower and higher temperature compressor bleed air from manifolds 48, 50 respectively and directs it through line 58 to the high pressure turbine casing 36. Control air may also be directed through line 60 to the low pressure turbine casing 38. <IMAGE>

Description

SPECIFICATION Gas turbine engine active clearance control Technical Field This invention relates to gas turbine engines, and more specifically to the active control of clearances between opposing seal elements of the rotor and stator assemblies.

Background Art It is well known in the gas turbine industry that engine performance is proportional to the leakage of working medium gases between opposing seal elements of the rotor and stator assemblies.

Techniques and concepts for reducing such clearances are continually under investigation and development.

One class of techniques are those relating to "active clearance control" in which the clearances are set as a function of engine operating condition. The objective is to establish minimum clearances under stable operating conditions, yet to provide sufficient clearance during transient operation to preclude destructive interference between relatively rotating components.

U.S. Patent Nos. 3,039,737 to Kolthoff entitled "Device for Controlling Clearance Between Rotor and Shroud of a Turbine"; 3,966,354 to Patterson entitled "Thermal Actuated Valve for Clearance Control": 3,975,901 to Hollinger et al entitled "Device for Regulating Turbine Blade Tip Clearance"; and 4,213,296 to Schwarz entitled "Seal Clearance Control System for a Gas Turbine" are representative of concepts and structures for effecting local control over rotor blade tip clearances. In some embodiments relatively hot air is utilized to move the seals away from the rotor blade tips and in other embodiments relatively cool air is utilized to move the seals toward the rotor blade tips. The concepts are at times combined in the same structure.

Recent commercial aircraft gas turbine engines, such as the JT9D-7R4 engine manufactured by Pratt S Whitney Aircraft, Division of United Technologies Corporation, have incorporated clearance control systems operative on a large segment of the engine to closely match thermal growth of the stator elements to that of the rotor elements. Principally, cooling or heating air is squirted onto the exterior of the engine case of the segment to be controlled. Desired contraction or expansion occurs.

U.S. Patent Nos. 4,069,662 to Redinger et al entitled "Clearance Control for Gas Turbine Engine"; 4,019,320 to Redinger et al entitled "External Gas Turbine Engine Cooling for Clearance Control"; and 4,279,123 to Griffin et al entitled "External Gas Turbine Engine Cooling for Clearance Control" are representative of the concepts employed in systems of the external type.

Advancing techniques for effecting segment cooling now include the wide distribution of cooling air at the interior of the case. Cooling air is flowed along the interior of the engine between the working medium flow path and the engine case.

U.S. Patent Nos. 3,957,391 to Vollinger entitled "Turbine Cooling"; 3,975,112 to Brown et al entitled "Apparatus for Sealing Gas Turbine Flow Path"; 4,005,946 to Brown et al entitled "Method anc Apparatus for Controlling Stator Thermal Growth"; and 4,242,042 to Schwarz entitled "Temperature Control of Engine Case for Clearance Control" representatively illustrate such concepts.

Notwithstanding the effectiveness of such prior art systems, scientists and engineers in the gas turbine engine industry are seeking yet improved systems employing judicious use of cooling/heating air.

Disclosure of the Invention According to the present invention the flow rate and temperature of turbine case, temperature modifying air in an active clearance control system is varied by modulating proportions of relatively low temperature, low pressure air and relatively high temperature, high pressure air in response to engine operation conditions.

In accordance with one detailed embodiment of the invention the case temperature modifying air is flowabie to one or more annular spaces circumscribing the cases to be controlled, and thence internally of the cases for cooling of components in proximity to the engine flow path.

A primary feature of the present invention is the utilization of dual source air for modifying the temperature of the engine case. Relatively low pressure, low temperature compressor air is mixed with relatively high temperature high pressure air at one or more modulating valves. The valves are capable of varying the proportions of air from each source for effecting case cooling at differing flow rates and temperatures.

In one detailed embodiment a shroud circumscribes each engine case to be controlled and is spaced apart therefrom. Case temperature modifying air is flowable to the space. The modifying air is subsequently flowable through apertures in the case into the interior of the engine for cooling components adjacent the engine flow path.

A principal advantage of the present invention is the judicious use of case temperature modifying air for controlling case diameter. Internal clearances at seals between rotor and stator structure are minimized by matching the case diameter to expected rotor growth under varied engine operating conditions. As viewed from another aspect, turbine cooling air utilized to protect engine components adjacent the flow path is diverted en route to preliminarily modify the temperature of the engine case.

Improved engine performance results from the sequential use of compressor air for such auxiliary purposes as well as from actual clearance control.

The foregoing features and advantages of the present invention will become more apparent in the light of the following detailed description of the best mode for carrying out the invention and in the accompanying drawing.

Brief Description of the Drawing Fig. 1 is a simplified side elevation view of a gas turbine engine with portions broken away in cross section; Fig. 2 is a simplified side elevation view of a portion of the engine illustrating the dual course of turbine case, temperature modifying air; Fig. 3 is a simplified view of a portion of the turbine section of the engine illustrating the distribution of cooling air internally of the engine; and Fig. 4 is a "pinch point" diagram illustrating relative thermal growth between the rotor and stator of such an engine.

Best Mode for Carrying Out the Invention An aircraft-type gas turbine engine capable of employing the concepts of the present invention is illustrated in the Fig. 1 partial cross section view. The engine principally includes a low pressure compression section 10, a high pressure compression section 12, a combustion section 14, a high pressure turbine section 1 6 and a low pressure turbine section 1 8. The engine illustrated is of the dual rotor type having a first shaft 20 joining a high turbine rotor assembly 22 to a high compressor rotor assembly 24 and a second shaft 26 joining a low turbine rotor assembly 28 to a low compressor rotor assembly 30.

The respective rotor assemblies are contained within a low compressor case 32, a high compressor case 34, a high turbine case 36 and a low turbine case 38. Rows of rotor blades, as represented by the single blades 40 extend outwardly on the rotor blades toward the engine cases.

Rows of stator vanes, as represented by the single vanes 42, are supported from the engine cases and extend inwardly therefrom in interdigitated position with respect to the blades 40. A flow path 44 from working medium gases extends axially through the engine between rows of rotor blades and rows of stator vanes.

The rows of rotor blades 40 are circumscribed by essentialiy cylindrical outer air seals 46. The positions of the outer air seals relative to the tips of the rotor blades is largely a function of the diameter of the engine case supporting the seals and of the temperature of the rotor blades. Particularly, within the turbine section the relative positions, referred to as "clearance" may vary widely over the operating range of the engine as the rotor blades and the case are subjected to differing thermal environments.

Curve A of Fig. 4 represents the radial position of the rotor blade tips at a turbine section location as a function of engine operating condition. Curve B of Fig. 4 represents the radial position of the outer air seal at the corresponding turbine location as a function of engine operating condition. The gap X between the two curves illustrates the expected clearance between the two relatively rotating components in an engine not employing the active clearance control concepts to be later described.

The simplified side elevation view of Fig. 2 illustrates apparatus incorporating the concepts of the present invention. A first manifold 48 is in gas communication with the compressor at a relatively low pressure, low temperature stage. A second manifold 50 is in gas communication with the compressor at a relatively high pressure, high temperature stage such as downstream of the final compression stage. A low pressure conduit 52 connects the manifold 48 with a modulating and mixing valve 54: a high pressure conduit 56 connects the manifold 50 to the valve 54.

The modulating and mixing valve 54 is capable for receiving the dual source air from the compressor and modulating the flow of each to produce an effluent having a desired temperature, pressure, and flow rate. In some embodiments the valve may be collaterally capable of producing dual effluents, each having individualized temperatures, pressures and flow rates. Effluent from the valve is flowed to the turbine section of the engine through one or more conduits 58. In the structure illustrated a second modulating and mixing valve on the reverse side of the engine is capable of discharging effluent through a second conduit 60 to a downstream position on the turbine. The first conduit 58 illustrated is capable of discharging to the high pressure turbine 16; the second conduit 60 illustrated is capable of discharging to the low pressure turbine 1 8.

The Fig. 3 turbine cross section view illustrates the distribution of effluent from the modulating and mixing valves via the first conduit 58 to the high pressure turbine 1 6 and via the second conduit 60 to the low pressure turbine 1 8. In the low turbine the case 38 is formed of double wall construction including an inner case 62 and an outer case or shroud 64. Effluent from the modulating and mixing valve is flowable to a space 66 between the inner case and shroud for the purpose of modifying the temperature of the case as a function of engine operating condition. The modifying air is hence flowable through apertures 68 in the inner case to the interior of the engine for subsequently cooling engine components in the turbine.

During operation of the engine, working medium gases are compressed within the compressor section to pressure ratios on the order of thirty to one (30:1 > and burned with fuel in the combustion section. The hot effluent from the combustion section is expanded through the turbine section to provide the motive force driving the compressor. Pressures across the compressor section of a typical engine increases at each succession stage from atmosphere pressure to the order of 31 bar at sea level take-off conditions. Correspondingly, temperatures acrosse the compressor section increase at each successive stage from ambient conditions to the order of 6200C at sea level take-off conditions.

Corresponding temperatures at the inlet to the turbine section are on the order of 13700 C. Radical variations in engine temperatures over the operating cycle of the engine establish the need for control of clearances between rotating and stationary structures under the influence of differing environments.

The concepts of the present invention employ case heating and case cooling in accord with the engine cycle to achieve close growth correspondence between the rotor and the case supported seals.

Case temperature modifying air is utilized for such heating and cooling. The modifying air comprises varied proportions of heating and cooling air ducted from the engine compressor to the case segment to be cooled. Representative characteristics of case temperature modifying air produced as the effluent from a modulating and mixing valve, such as that described herein, is shown in the table reproduced below. The pressure, temperature and flow rate data is representative of a 1 8000 kp thrust class engine at idle, sea level takeoff and cruise conditions. Data is for a split-type system in which a first modulating valve is supplied with dual source air for discharge and temperature control of the high pressure turbine case and a second modulating valve is supplied with dual source air for discharge and temperature control of the low pressure turbine case.

HIGH PRESSURE TURBINE Low Pressure High Pressure Low Temperature High Temperature High Turbine Source Source Modifying Air Pressure 1,86 bar 4,2 bar 1,72 bar Idle Temp 143"C 221"C 221"C Flow Rate 0.0 kg/sec 0,027 kg/sec 0,027 kg/sec Sea Pressure 9,37 bar 29,9 bar 8,96 bar Level Temp 382"C 599or 521"C Takeoff Flow Rate 0;;045 kg/sec 0,099 kg/sec 0,145 kg/sec Pressure 4,48 bar 13,58 bar 4,13 bar Cruise Temp 304or 482'C 304"C Flow Rate 0,070 kg/sec 0,0 kg/sec 0,070 kgisec LOW PRESSURE TURBINE Low Pressure High Pressure Low Temperature High Temperature High Turbine Source Source Modifying Air Pressure 1,37 bar 4,2 bar -0- Idle Temp 104"C 221"C -0 Flow Rate 0,0 kg/sec 0;0 kg/sec -0- s I .

Sea Pressure 5,17 bar 29,7 bar 4,27 bar Level Temp 304"C 599"C 438 "C Takeoff Flow Rate 0,29 kg/sec 0,238 kg/sec 0,529 kg/sec I Pressure 2,27 bar 13,58 bar 1,93 bar Cruise Temp 216"C 482"C 216 0C Flow Rate 0,254 kg/sec 0,0 kg/sec 0,254 kg/sec Each of the one or more modulating valves is controllable in response to engine operating conditions to produce the effluents described above. The modulating valves are controllable in response to engine operating conditions. Parameters respresentative of engine condition, such as case temperature, rotor speed, engine pressure rates, altitude Mach Number, turbine temperature and exhaust gas temperature are selected for control.For the representative engine described above the parameters shaft RPM, altitude, and flight Mach Number were selected for control.

Low Rotor High Rotor Flight Speed Speed Altitude Mach Number Ground Idle 1115 RPM 10,063 RPM 0 m 0.0 Sea Level Take Off 3923 RPM 14,045 RPM 0 m 0.0 Cruise 3902 RPM 13,178 RPM 10668 m .80 Referring again to the Fig. 4 "pinch point" diagram curve C represents the radial position of an outer air seal as it is varied over the engine operating range by modifying the supporting case in accordance with the present concepts in accordance with the sensed parameters, shaft RPM, altitude and flight Mach Number. The gap Y represents the attainable relative clearance between the tips of the rotor blades and the corresponding outer air seal. Clearance is not only greatly reduced from the noncontrolled conditions, but closely corresponds in contour to the radial position of the tips. The minimum clearance necessary to avoid destructive interferences is provided.

Although the invention has been described with respect to a particular turbine embodiment, it should be understood that the invention is not so limited and that various changes and modifications may be made without departing from the spirit and scope of this novel concept.

Claims (8)

1. A method for controlling the clearance between opposing seal elements of the rotor assembly and the stator assembly of a gas turbine engine comprising the steps of: flowing relatively low pressure, low temperature air from the compressor of the engine to a modulating and mixing valve; flowing relatively high pressure, high temperature air from the compressor of the engine to said modulating and mixing valve; mixing said relatively low pressure, low temperature air and said relatively high pressure, high temperature air at the modulating valve in proportions functionally related to engine operating condition; and flowing said mixed air to the turbine section of the engine and against the case thereof for thermally varying the diameter of the case to achieve control over clearances between the rotor and stator assemblies.

2. The method according to claim 1 wherein the proportions of relatively low pressure, low temperature air and relatively high pressure, high temperature air are selected to produce a mixture of air having a desired temperature, pressure and flow rate for judicious use in thermally modifying the diameter of the turbine case emanating from the modulating valve.

3. The method according to claim 1 which further includes the steps of: flowing relatively low pressure, low temperature air from the compressor of the engine to a second modulating and mixing valve; flowing relatively high pressure, high temperature air from the compressor of the engine to said second modulating and mixing valve; mixing said relatively low pressure, low temperature air and said relatively high pressure, high temperature air at the second modulating valve in proportions functionally related to engine operating condition; and flowing said air mixed at the second modulating valve to the turbine section at a location downstream of the location to which the air mixed at the first modulating valve was flowed and against the case at that downstream location for thermally varying the diameter of the case at that location.

4. The method according to claim 1 wherein the mixture of air flowing to the turbine at sea level takeoff condition comprises both air from the low pressure, low temperature and high pressure, high temperature portions and wherein the mixture at cruise condition comprises air solely from the low pressure, low temperature portion of the compressor.

5. The method according to claim 1, 2, 3 or 4 wherein the proportions of low pressure, low temperature and high pressure, high temperature air flowed to and mixed at the modulating valve air varied in response to one or more parameters selected from the group consisting of low rotor speed, high rotor speed, altitude and flight Mach number.

6. In a gas turbine engine of the type having compression and turbine sections formed of a rotor assembly circumscribed by a stator assembly and including means for varying the diameter of the turbine case to control the clearance between the rotor and stator assemblies, the improvement comprising:: a modulating air mixing valve capable of receiving relatively low pressure, low temperature and relatively high pressure, high temperature air from the compressor, mixing said air in proportions functionally related to engine operating condition and discharging the resultant mixture; a manifold at the compressor for collecting relatively low pressure, low temperature air; a conduit joining the manifold for collecting relatively low pressure, low temperature air to the modulating valve; a manifold at the compressor for collecting relatively high pressure, high temperature air; a conduit joining the manifold for collecting relatively low pressure, low temperature air to the modulating valve; a conduit for flowing the air mixture discharged from the modulating valve to the turbine case.

7. The invention according to claim 6 wherein the turbine case is of a double wall construction having an inner case and shroud spaced apart from the inner case and wherein said conduit for flowing the air mixture is capable of discharging said mixture into the space between the inner case and the shroud for modifying the temperature of the case.

8. The invention according to claim 7 wherein said inner case has a plurality of apertures therethrough for enabling air discharge to the space between the inner case and the shroud to subsequently flow to the interior of the case for cooling components of the turbine.