公开(公告)号：JP2000186508A

公开(公告)日：2000-07-04

申请号：JP33761799

申请日：1999-11-29

Applicant: UNITED TECHNOLOGIES CORP

Inventor： IRWIN CRAIG W ,  KHALID SYED J

IPC: F01D11/08 ,  F01D11/24 ,  F01D25/00 ,  F02C7/00 ,  F02C7/28 ,  F02C9/28

Abstract: PROBLEM TO BE SOLVED: To specify the engine operating condition and a thermal condition determined by the condition without determining a steady state temperature by giving a signal showing a difference between the instant clearance under the thermal condition and the steady state clearance in the engine operating condition in response to a signal showing the engine operating condition. SOLUTION: A gas turbine engine 20 includes a fan area stored in an outer engine case 32 along a longitudinal axis 30, a compression area 24, a combustion area 26 and a turbine area 28. The gas turbine engine 20 has a passage 34 where working medium gas flows. The gas passage 34 is extended through the respective stream areas. The compression area 24 has a rotor assembly 40 provided in a case 44 and a stator assembly 42. A blade 54 is arranged in the circumferential direction round a rim 58 and fitted to the rim 58. The blade 54 and the case 44 are provided at a space to define a clearance.

公开(公告)号：JP2000179360A

公开(公告)日：2000-06-27

申请号：JP35250499

申请日：1999-12-13

Applicant: UNITED TECHNOLOGIES CORP

Inventor： IRWIN CRAIG W ,  KHALID SYED J

IPC: F02C9/00 ,  F04D27/02

Abstract: PROBLEM TO BE SOLVED: To enable increase of the stall margin through control of bleed valve by calculating the rate of temperature change of the metallic materials of the compressor based on the detected values of the temperature and pressure of the gas flowing within the compressor and obtaining the signal which indicates the degree of compressor instability due to the effect of thermal transmission. SOLUTION: In a double-shaft gas turbine engine, a temperature sensor 136 and a pressure sensor 138 are located in the gas passage at the compressor exit. The signals detected by the sensors are sent to a metallic materials temperature calculation logical block 160 wherein the primary delay of the gas flow temperature is added to calculate the temperature proportionate to the metallic materials temperature of the casing and stator/blades. Then, thermal transmission parameter is obtained by differentiating the metallic materials temperature at a thermal transmission calculation logical block 168. Next, nondimensional thermal transmission parameter is obtained by means of a logical standardization block 176 by dividing the thermal transmission parameter by the product of the gas flow rate and the gas temperature. Thus obtained nondimensional thermal transmission parameter is then compared with the threshold value calculated by means of a threshold value logical block 184, and a bleed valve 148 of the compressor is controlled accordingly.

公开(公告)号：US6272422B2

公开(公告)日：2001-08-07

申请号：US75359401

申请日：2001-01-04

Applicant: UNITED TECHNOLOGIES CORP

Inventor： KHALID SYED J ,  IRWIN CRAIG W

IPC: F01D11/14 ,  F01D21/00 ,  F01D21/04 ,  F02C9/00

CPC classification number: F01D11/14 ,  F01D21/003 ,  F01D21/04

Abstract: A method and an apparatus for determining the clearance between the rotor blades of a rotor assembly and a shroud disposed radially outside of the rotor assembly is provided that calculates steady-state operating conditions for a given power engine setting and utilizes those steady-state conditions to determine a steady-state clearance at the given power setting. The method and apparatus further calculate instantaneous thermal conditions for the rotor disk, rotor blades, and shroud. The instantaneous thermal conditions are subsequently used to determine the amount of instantaneous thermal expansion of the rotor disk, rotor blades, and shroud. A clearance transient overshoot is determined using the calculated instantaneous thermal expansions. The actual clearance is determined using the steady-state clearance and the clearance transient overshoot.

Abstract translation: 提供了一种用于确定转子组件的转子叶片和设置在转子组件的径向外侧的护罩之间的间隙的方法和装置，其计算给定动力发动机设定的稳态操作条件，并将这些稳态条件 在给定功率设置下确定稳态间隙。 该方法和装置进一步计算转子盘，转子叶片和护罩的瞬时热条件。 瞬时热条件随后用于确定转子盘，转子叶片和护罩的瞬时热膨胀量。 使用计算的瞬时热膨胀确定清除瞬态过冲。 实际间隙使用稳态间隙和间隙瞬变过冲来确定。

公开(公告)号：DE69924959T2

公开(公告)日：2005-09-29

申请号：DE69924959

申请日：1999-12-16

Applicant: UNITED TECHNOLOGIES CORP

Inventor： IRWIN CRAIG W ,  KHALID SYED J

IPC: F01D11/08 ,  F01D11/24 ,  F01D25/00 ,  F02C7/00 ,  F02C7/28 ,  F02C9/28 ,  F01D11/14 ,  F01D11/16 ,  F01D11/18

Abstract: A method and an apparatus for use with a gas turbine engine (20) receive a signal indicative of an engine operating condition, generate signals representative of thermal conditions of a rotor (52), blades (54), and a case (44), and in response to each of the signals above, determine a signal indicative of a difference between a instantaneous clearance (60) for the thermal conditions and a steady state clearance for the engine operating condition. The determination includes effects related to a temporary difference that results from a difference between the steady state clearance for the engine operating condition and a steady state clearance for a preceding engine operating condition, but does not require computation of the actual temperatures or the steady state temperatures of the rotor (52), the blades (54), and the case (44). A signal indicative of the difference between a instantaneous clearance for the thermal conditions and a steady state clearance (60) for the engine operating condition may be provided to various augmentation schedules.

公开(公告)号：DE68928159T2

公开(公告)日：1997-10-30

申请号：DE68928159

申请日：1989-09-13

Applicant: UNITED TECHNOLOGIES CORP

Inventor： POLLAK ROBERT R ,  MARCOS JUAN A ,  KHALID SYED J

IPC: F02C9/50 ,  F02C9/28 ,  F02C9/54 ,  F02K1/16 ,  F02K1/17

Abstract: An active control system for use in gas turbine engines synchronizes exhaust nozzle (26) area and burner fuel flow together with gas path variable engine parameters, such as fan variable vane and high compressor variable vane positions (36, 38, 40). As a result, extremely fast thrust transients are possible with optimized compression system stability, since fan and compressor rotor speeds are held high, allowing total engine power to be controlled by air flow and fuel flows directly. Time responsiveness to thrust demands for a fan jet engine of the type that includes twin spools and a variable inlet at the fan is enhanced by regulating air flow through the variable inlet while holding low pressure compressor speed constant. The thrust change is targeted as a function of power lever position which generates a low pressure compressor parametric as a function of engine and aircraft operating variables. This parameter is then utilized to control both the inlet variable vanes and fuel flow to 1) attain the targeted value and 2) return the engine to its steady state operating line after a given time interval. The time responsiveness and engine stability for acceleration and deceleration engine transient is enhanced by control mechanism that synchronously adjusts fuel flow and high pressure compressor vane position as a function of corrected low pressure compressor speed in a twin spool axial flow turbine power plant. The acceleration and deceleration mode is targeted by power lever position and the target is attained by adjusting the vane position at a constant high pressure compressor speed and returned to a fuel efficient and engine stable steady state operating line by concomittantly adjusting fuel flow and vane position.

公开(公告)号：DE69925231T2

公开(公告)日：2005-10-20

申请号：DE69925231

申请日：1999-12-10

Applicant: UNITED TECHNOLOGIES CORP

Inventor： IRWIN CRAIG W ,  KHALID SYED J

IPC: F02C9/00 ,  F04D27/02

Abstract: A control system (104) for controlling the compressor stall margin during acceleration in a gas turbine engine (100) includes means (132,138) for sensing signals indicative of the gas flow temperature and gas pressure. The control system (104) further includes signal processing means (144), responsive to the sensed signals, for synthesizing and providing a processed signal indicative of a measure of compressor destabilization due to heat transfer effects. The control system (104) also includes means (152), responsive to the processed signal, for providing an output (156) to initiate corrective action to increase compressor stall margin if needed. The engine control means (152), which is a part of the control system (104), increases the compressor stall margin by either adjusting the compressor variable vanes, reducing fuel flow or modulating the compressor bleed.

公开(公告)号：DE69925231D1

公开(公告)日：2005-06-16

申请号：DE69925231

申请日：1999-12-10

Applicant: UNITED TECHNOLOGIES CORP

Inventor： IRWIN CRAIG W ,  KHALID SYED J

IPC: F02C9/00 ,  F04D27/02

Abstract: A control system (104) for controlling the compressor stall margin during acceleration in a gas turbine engine (100) includes means (132,138) for sensing signals indicative of the gas flow temperature and gas pressure. The control system (104) further includes signal processing means (144), responsive to the sensed signals, for synthesizing and providing a processed signal indicative of a measure of compressor destabilization due to heat transfer effects. The control system (104) also includes means (152), responsive to the processed signal, for providing an output (156) to initiate corrective action to increase compressor stall margin if needed. The engine control means (152), which is a part of the control system (104), increases the compressor stall margin by either adjusting the compressor variable vanes, reducing fuel flow or modulating the compressor bleed.

公开(公告)号：DE69515814D1

公开(公告)日：2000-04-27

申请号：DE69515814

申请日：1995-12-05

Applicant: UNITED TECHNOLOGIES CORP

Inventor： KHALID SYED J

IPC: F02C7/28 ,  F01D11/08 ,  F01D11/12 ,  F04D27/02

Abstract: The compressor section of a gas turbine engine contains a insert (24) installed around the compressor blades that includes cells (28) in a honeycomb configuration. Each cell is at a compound angle to the blade tip to energize the tip air flow as the tip passes over the cell as the blade rotates, improving the stall margin. Each cell is oriented in the direction of the blade chord and facing the advancing blades. As the blade rotates it sweeps by each cell and high pressure airflow is first captured in the cell from the high pressure side of the blade and released to the low pressure side as the blade passes the cell, creating a energizing jet of high velocity flow in the direction of the airflow across the blade.

公开(公告)号：DE69924959D1

公开(公告)日：2005-06-02

申请号：DE69924959

申请日：1999-12-16

Applicant: UNITED TECHNOLOGIES CORP

Inventor： IRWIN CRAIG W ,  KHALID SYED J

IPC: F01D11/08 ,  F01D11/24 ,  F01D25/00 ,  F02C7/00 ,  F02C7/28 ,  F02C9/28 ,  F01D11/14 ,  F01D11/16 ,  F01D11/18

Abstract: A method and an apparatus for use with a gas turbine engine (20) receive a signal indicative of an engine operating condition, generate signals representative of thermal conditions of a rotor (52), blades (54), and a case (44), and in response to each of the signals above, determine a signal indicative of a difference between a instantaneous clearance (60) for the thermal conditions and a steady state clearance for the engine operating condition. The determination includes effects related to a temporary difference that results from a difference between the steady state clearance for the engine operating condition and a steady state clearance for a preceding engine operating condition, but does not require computation of the actual temperatures or the steady state temperatures of the rotor (52), the blades (54), and the case (44). A signal indicative of the difference between a instantaneous clearance for the thermal conditions and a steady state clearance (60) for the engine operating condition may be provided to various augmentation schedules.

公开(公告)号：DE69515814T2

公开(公告)日：2000-10-12

申请号：DE69515814

申请日：1995-12-05

Applicant: UNITED TECHNOLOGIES CORP

Inventor： KHALID SYED J

IPC: F02C7/28 ,  F01D11/08 ,  F01D11/12 ,  F04D27/02

Abstract: The compressor section of a gas turbine engine contains a insert (24) installed around the compressor blades that includes cells (28) in a honeycomb configuration. Each cell is at a compound angle to the blade tip to energize the tip air flow as the tip passes over the cell as the blade rotates, improving the stall margin. Each cell is oriented in the direction of the blade chord and facing the advancing blades. As the blade rotates it sweeps by each cell and high pressure airflow is first captured in the cell from the high pressure side of the blade and released to the low pressure side as the blade passes the cell, creating a energizing jet of high velocity flow in the direction of the airflow across the blade.