# US 12,000,290 B2 · Control method and unit for controlling the clearance of a high-pressure turbine to reduce the effect of EGT overshoot

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- Number: US 12,000,290 B2
- Document ID: US12000290B2
- Status: Granted
- Kind: B2
- Application number: 17/757,981
- Filed: 2021-01-04
- Published: 2023-02-09
- Granted: 2024-06-04
- Earliest claimed: 2020-01-08
- Assignees: Safran Aircraft Engines
- Applicants: Safran Aircraft Engines
- Inventors: Jérôme Claude George Lemonnier; Damien Bonneau; Cécile Racinet; Antoine Bruno Van Noort
- Web page: https://patentgrep.com/patent/US12000290B2

## CPC

- F01D 11/14: Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing (rotors with blades adjustable in operation F01D7/00)
- F01D 11/24: by selectively cooling-heating stator or rotor components
- F01D 19/02: dependent on temperature of component parts, e.g. of turbine-casing
- F01D 21/12: responsive to temperature
- F01D 25/12: Cooling
- F05D 2220/323: for aircraft propulsion, e.g. jet engines
- F05D 2240/55: Seals
- F05D 2260/201: by impingement of a fluid
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

Method for controlling a clearance between the tips of the blades of a rotor of an aircraft engine turbine and a turbine ring, comprising the estimation of the clearance to be controlled and the control of a valve delivering an air stream directed towards the turbine ring based on the thus estimated clearance, this method comprising: the detection of a transient acceleration phase based on at least one parameter representative of the engine; the receipt of a data relating to the altitude of the aircraft; the determination of data representative of the temperature of the rotor during the transient acceleration phase and in steady speed and the calculation of a relative temperature deviation.

## Claims

### Claim 1

A method for controlling a clearance between tips of blades of a rotor of a high-pressure turbine of a gas turbine aircraft engine of an aircraft and a turbine ring of a casing surrounding said blades of the high-pressure turbine, the method comprising:
detecting a transient phase of acceleration of the engine based on at least one parameter representative of the engine;
receiving a data relating to an altitude of the aircraft;
determining data representative of a temperature of said rotor of the high-pressure turbine of the engine during the transient acceleration phase subsequent to determining data representative of a temperature of said rotor of the high-pressure turbine of the engine during a steady speed;
calculating a relative temperature deviation between said temperature data of the transient acceleration phase and the temperature data of the steady speed;
estimating the clearance to be controlled;
if said transient acceleration phase is detected and if said relative temperature deviation is greater than a predetermined minimum temperature deviation, defining a level of opening and a time of opening of a valve by a predefined altitude/relative temperature deviation lookup table; and
controlling the opening of the valve at the opening level and during the opening time defined to deliver an air stream directed to the turbine ring based on said estimated clearance.

### Claim 2 (depends on 1)

The method according to claim 1, for which the level of opening and the time of opening of the valve are defined from a predefined altitude/relative temperature deviation lookup table only if said estimated clearance is also greater than a predetermined minimum clearance.

### Claim 3 (depends on 1)

The method according to claim 1, for which the control of the opening of the valve also includes a timing on its opening defining a time limit from which the valve is open at a desired opening level and for a desired opening time following the detection of the transient acceleration phase.

### Claim 4 (depends on 1)

The method according to claim 1, wherein the transient phase of acceleration of the engine is detected from the deviation between the temperature data of the steady speed (Tstab) and the temperature data of the transient acceleration phase (Ttrans).

### Claim 5 (depends on 1)

The method according to claim 1, wherein said at least one parameter representative of the engine is chosen among: a speed of a low-pressure turbine of the engine, a speed of the high-pressure turbine, a pressure measured at a high-pressure compressor, an angular position of a throttle control lever of the aircraft and the data representative of a temperature of gases at an outlet of a combustion chamber of the engine.

### Claim 6 (depends on 1)

The method according to claim 1, wherein the data representative of the temperature of the rotor during the transient acceleration phase or the steady speed is an estimate of a temperature of a rotor disk of the high-pressure turbine based on said at least one parameter representative of the engine.

### Claim 7

A control unit for controlling a clearance between tips of blades of a rotor of a high-pressure turbine of a gas turbine aircraft engine of an aircraft and a turbine ring of a casing surrounding said blades of the high-pressure turbine, the control unit comprising:
means for estimating the clearance to be controlled;
means for controlling a valve, the valve being configured to deliver an air stream towards said turbine ring based on said estimated clearance to be controlled;
detection means configured to detect a transient phase of acceleration of the engine based on at least one parameter representative of the engine;
receiving means configured to receive a data relating to an altitude of the aircraft;
calculation means configured to determine data representative of a temperature of said rotor of the high-pressure turbine of the engine during the transient acceleration phase subsequent to determining data representative of a temperature of said rotor of the high-pressure turbine of the engine during a steady speed, and to calculate a relative temperature deviation between said temperature data determined during the transient acceleration phase and the temperature data determined during the steady speed; and
control means configured to control an opening of the valve to deliver said air stream to the turbine ring at an opening level and for an opening time defined by a predefined altitude/relative temperature deviation lookup table, if the transient acceleration phase is detected and if said relative temperature deviation is greater than a predetermined minimum temperature deviation.

### Claim 8 (depends on 7)

The control unit according to claim 7, wherein the predefined altitude/relative temperature deviation lookup table delivers a value of the parameters {X, Y(i), Z(i)} for a given pair of values {Altitude; relative deviation (Tstab−Ttrans)/Tstab}, with: X defining a time limit after which the valve must open, Z(i) defining, for a given plateau i, the valve opening level, Y(i) defining the valve opening time for the plateau i and the opening level Z(i).

### Claim 9 (depends on 7)

The control unit according to claim 7, wherein the valve is a position-regulated valve intended to remain in a closed position when without a power supply[,] and whose position can be comprised between 0% (full closing), corresponding to a closed valve, and 100% (full opening).

### Claim 10 (depends on 7)

A gas turbine aircraft engine comprising a control unit according to claim 7 and at least one valve to act on an air stream directed towards the turbine ring and wherein the valve is controlled by the control means.