Frost protection system for an aircraft engine nacelle

Patent metadata

Field	Value
Publication number	US-11220343-B2
Application number	US-201816156272-A
Country	US
Kind code	B2
Filing date	Oct 10, 2018
Priority date	Oct 20, 2017
Publication date	Jan 11, 2022
Grant date	Jan 11, 2022

Abstract

A frost protection system for an aircraft engine nacelle, the nacelle comprising an inner shroud provided with at least one acoustic panel and an air intake lip forming a leading edge of the nacelle. The protection system comprises a heat exchanger device including at least one heat pipe configured to transfer heat emitted by a heat source to the at least one acoustic panel.

First claim

The invention claimed is:

1. An aircraft engine nacelle, comprising:
an inner shroud provided with at least one acoustic panel, and
a frost protection system comprising:
a heat exchanger device including at least one heat pipe configured to convey a heat transfer fluid and to transfer heat emitted by a heat source to the at least one acoustic panel,
at least one evaporator thermally connected to the heat source, the at least one evaporator being configured to extract at least a portion of the heat supplied by the heat source, the extracted heat being transferred to the heat transfer fluid;
at least one condenser attached to the inner shroud, the at least one condenser being configured to deliver at least a portion of the heat extracted by the at least one evaporator to the at least one acoustic panel, the extracted heat being transferred to the at least one condenser via the heat transfer fluid;
each of the at least one evaporators being fluidically connected to at least one of the at least one condenser by at least one heat pipe through which the heat transfer fluid flows, and
the nacelle comprises an air intake lip forming a leading edge of the nacelle, the lip having an annular space, the annular space being closed by an internal partition and being arranged to receive a hot air supply,
wherein the heat source corresponds to the internal partition,
wherein the at least one evaporator is connected by attachment to the internal partition, the at least one evaporator is configured to extract at least a portion of the heat supplied through the internal partition by hot air being supplied to the annular space of the lip, the extracted heat being transferred to the heat transfer fluid,
a source of the hot air being supplied to the annular space of the nacelle.

2. The aircraft engine nacelle according to claim 1 , wherein the at least one heat pipe comprises:
at least one steam pipe configured to convey, from the at least one evaporator to the at least one condenser, the heat transfer fluid vaporized by the heat extracted by the at least one evaporator;
at least one liquid pipe configured to convey, from the at least one condenser to the at least one evaporator, the heat transfer fluid condensed by cooling in the condenser.

3. The aircraft engine nacelle according to claim 2 , comprising:
an air intake lip forming a leading edge of the nacelle, the lip having an annular space, the annular space being closed by an internal partition and being arranged to receive a hot air supply, the at least one evaporator being connected by attachment to the internal partition, the at least one evaporator being configured to extract at least a portion of the heat supplied through the internal partition by hot air being supplied to the annular space of the lip, the extracted heat being transferred to the heat transfer fluid.

4. The aircraft engine nacelle according to claim 1 , wherein the at least one condenser comprises at least one heating channel.

5. An aircraft fitted with at least one engine and comprising according to claim 1 , the least one engine being surrounded by the nacelle.

6. The aircraft according to claim 5 , further comprising:
at least one air-heating device configured to produce the hot air being supplied to the annular space of the nacelle,
at least one duct linking the at least one air-heating device to the annular space of each nacelle, the at least one duct being configured to convey the hot air produced by the air-heating device to the annular space of the lip,
at least one valve for each of the at least one duct which is configured to regulate a pressure and a flow rate of the hot air flowing through the at least one duct.

7. The aircraft according to claim 5 , wherein the air-heating device corresponds to compression stages of the engine surrounded by the nacelle.

8. The aircraft according to claim 5 , further comprising an electrical system corresponding to the heat source.

9. The aircraft according to claim 8 , wherein the electrical system corresponds to an electrical device dedicated to production of heat for the frost protection system.

10. The aircraft according to claim 8 , wherein the electrical system supplies electrical power to the aircraft.

Assignees

• Airbus Operations Sas

Inventors

• Porte Alain
• Carcone Jonathan

Classifications

• F05D2260/208

  using heat pipes · CPC title

• B64D15/04 Primary

  Hot gas application · CPC title

• B64D2033/0233

  comprising de-icing means · CPC title

• F02C7/045

  having provisions for noise suppression · CPC title

• B64D33/02

  of combustion air intakes · CPC title