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US-12618377-B2 - Unducted propeller gas turbine comprising a cooling air channel and a variable bleed valve ejection channel

US12618377B2US 12618377 B2US12618377 B2US 12618377B2US-12618377-B2

Abstract

A gas turbine includes two unducted propellers, a main channel, a compression stage located in the main channel, the main channel supplying the compression stage with air, a cooling channel extending from a cooling inlet opening into the main channel, the cooling inlet being located upstream of the compression stage with reference to a direction of flow of the air through the gas turbine, a heat exchanger located in the cooling channel, an ejection channel opening into the main channel and into the cooling channel so that air passing through the ejection channel ventilates the heat exchanger, and a variable bleed valve that adjust an air flow rate in the ejection channel.

Inventors

  • Tom EVERAETS

Assignees

  • SAFRAN AIRCRAFT ENGINES
  • GENERAL ELECTRIC COMPANY

Dates

Publication Date
20260505
Application Date
20220404

Claims (14)

  1. 1 . A method for controlling a gas turbine, the method comprising: supplying air into a main channel through a main inlet, the main inlet being located between two unducted propellers of the gas turbine, producing compressed air by flowing the air supplied through a compression stage located in the main channel, producing a complementary airflow through an ejection channel, the ejection channel opening into the main channel and into a cooling channel, so that air passing through the ejection channel ventilates ana heat exchanger located in the cooling channel, the cooling channel extending from a cooling inlet, the cooling inlet opening into the main channel, the cooling inlet being located upstream of the compression stage with reference to a direction of flow of the air through the gas turbine, comparing a rotational speed of one of the unducted propellers with a first predetermined speed, and if the rotational speed exceeds the first predetermined speed, setting to zero a flow of the complementary airflow.
  2. 2 . The method according to claim 1 , wherein the first predetermined speed is equal to 80% of a maximum rotational speed of the unducted propeller.
  3. 3 . The method according to claim 1 , further comprising producing a cooling airflow flowing from the main channel into the cooling channel.
  4. 4 . The method according to claim 3 , further comprising controlling a flow of the cooling airflow based on a cooling requirement of the heat exchanger.
  5. 5 . A gas turbine comprising: two unducted propellers, a main channel comprising a main inlet, the main inlet being located between the two unducted propellers, a compression stage located in the main channel, the main channel being configured to supply the compression stage with air, a cooling channel extending from a cooling inlet opening into the main channel, the cooling inlet being located upstream of the compression stage with reference to a direction of flow of the air through the gas turbine, a heat exchanger located in the cooling channel, an ejection channel opening into the main channel and into the cooling channel so that air passing through the ejection channel ventilates the heat exchanger, and a variable bleed valve configured to adjust an air flow rate in the ejection channel.
  6. 6 . The gas turbine according to claim 5 , wherein the variable bleed valve comprises a rotary valve configured to be rotated.
  7. 7 . The gas turbine according to claim 5 , wherein the variable bleed valve comprises a sliding door.
  8. 8 . The gas turbine according to claim 5 , further comprising an airflow modulator configured to adjust a flow rate of an airflow passing through the cooling inlet from the main channel to the cooling channel.
  9. 9 . The gas turbine according to claim 8 , wherein the airflow modulator is placed at the cooling inlet.
  10. 10 . The gas turbine according to claim 8 , wherein the airflow modulator is placed at a cooling outlet of the cooling channel, the modulator comprising a rotary valve configured to be rotated.
  11. 11 . The gas turbine according to claim 8 , wherein the airflow modulator comprises a variable section of the cooling channel.
  12. 12 . An aircraft comprising the gas turbine according to claim 5 .
  13. 13 . The gas turbine according to claim 5 , wherein the heat exchanger is located at an intersection between the cooling channel and the ejection channel.
  14. 14 . The gas turbine according to claim 5 , further comprising a speed reduction gearbox configured to drive in rotation one of the unducted propellers, the heat exchanger being configured to ensure a cooling of the speed reduction gearbox.

Description

CROSS REFERENCE TO RELATED APPLICATION This application is a National Stage of International Application No. PCT/FR2022/050626 filed Apr. 4, 2022, the contents of which being herein incorporated by reference in its entirety. FIELD OF THE INVENTION The invention relates to a propulsion gas turbine with an unducted propeller and particularly the air cooling in such systems. The invention is of particular interest when it is applied to gas turbomachines for aircraft propulsion. STATE OF THE ART It is known a gas turbine 10 comprising a fan composed of an unducted upstream propeller and, downstream, of an unducted downstream propeller or fixed rectifier. The terms upstream and downstream are defined in relation to the general direction of flow of the gases through the gas turbine. The propulsion system extends along an axis and includes successively, in the direction of flow of the gases in the turbomachine, the fan, an air inlet configured to supply with air: a compression section which may comprise a low-pressure compressor and a high-pressure compressor,a combustion chamber,a turbine section which may comprise a high-pressure turbine, anda low-pressure turbine. The air entering through the air inlet is finally expelled from the gas turbine through a main outlet, located downstream of the low-pressure turbine. The upstream propeller and the low-pressure compressor are driven in rotation by the low-pressure turbine via a first transmission shaft, while the high-pressure compressor is driven in rotation by the high-pressure turbine via a second transmission shaft. In operation, an air flow is compressed by the low-pressure and high-pressure compressors and supplies combustion into the combustion chamber, the expansion of the combustion gases of which drives the high-pressure and low-pressure turbines. The air propelled by the upstream propeller and the combustion gases exiting through the gas outlet downstream of the turbines exert a reaction thrust on the propulsion system and, through it, on a vehicle or machine such as an aircraft. A speed reduction gearbox is driven by the first transmission shaft, and then drives the upstream propeller which rotates at a reduced rotational speed compared to the rotational speed of the transmission shaft. The air entering through the air inlet flows through a channel up to the low-pressure compressor. Upstream of the low-pressure compressor, a rotor with moving blades which is placed in the channel is driven by the transmission shaft. The rotor with moving blades comprises a plurality of compression blades and increases, downstream, the pressure of the air circulating in the channel. A stator with fixed vanes is located upstream and/or downstream of the rotor with moving blade. A complementary channel extends from the channel, downstream of the input rectifier, up to a complementary outlet opening onto the outside of the body of the propulsion system. The complementary channel has an annular shape and extends about the axis of the propulsion system. The air outlet is located downstream of the downstream fixed propeller and upstream of the main outlet. The complementary channel allows cooling a heat exchanger, in particular with the aim of cooling the speed reduction gearbox. Such a complementary channel also contributes to part of the thrust of the engine. However, such a complementary channel complicates the structure of the propulsion system. There is therefore a need for a simpler structure of the complementary channel which ensures a cooling function. DISCLOSURE OF THE INVENTION One aim of the invention is to propose a simpler structure allowing greater cooling power than in the prior art. The aim is achieved within the framework of the present invention thanks to a gas turbine comprising: two unducted propellers,a main channel,at least one compression stage located in the main channel, the main channel being configured to supply the compression stage with air,a cooling channel extending from a cooling inlet opening into the main channel, the cooling inlet being located upstream of the or each compression stage with reference to a direction of flow of the air,a heat exchanger located in the cooling channel,an ejection channel opening into the main channel and into the cooling channel so that the air passing through the ejection channel ventilates the heat exchanger, anda variable bleed valve configured to adjust an air flow rate in the ejection channel. The terms upstream and downstream are defined in relation to the direction of flow of the air through the gas turbine. The cooling inlet being located upstream of any compression stage in the main channel, all the compression stages located in the main channel are thus grouped together downstream of the cooling inlet, which allows simplifying the structure relative to the prior art. In addition, the air entering the cooling channel is not, unlike the prior art, heated by a compression stage located in the main channel up