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CN-122029038-A - Component comprising a fire barrier

CN122029038ACN 122029038 ACN122029038 ACN 122029038ACN-122029038-A

Abstract

The invention relates to a component (11) for an aircraft turbomachine, the component (11) comprising-a body (12) comprising a composite material comprising an organic matrix and fibres embedded in the matrix, and-a fire protection barrier (13) arranged on at least one surface of the body (12), characterized in that the fire protection barrier (13) is multi-layered and comprises-a perforated first layer (14) comprising glass fibres, and-a second layer (15) comprising fibres other than glass fibres, the first layer (14) being located between the body (12) and the second layer (15).

Inventors

  • MATHON RICHARD

Assignees

  • 赛峰飞机发动机公司

Dates

Publication Date
20260512
Application Date
20241010
Priority Date
20231018

Claims (10)

  1. 1. A component (11) for an aircraft turbomachine (1), the component (11) comprising: -a body (12) comprising a composite material comprising an organic matrix and fibres embedded in said matrix, and -A fire protection barrier (13) arranged on at least one surface of the body (12), characterized in that the fire protection barrier (13) is multi-layered and comprises: -a perforated first layer (14), said perforated first layer comprising glass fibres, and A second layer (15) comprising fibres other than glass fibres, The first layer (14) is located between the body (12) and the second layer (15).
  2. 2. The component according to the preceding claim, characterized in that the second layer (15) comprises carbon fibers.
  3. 3. The component according to any one of the preceding claims, wherein the first layer (14) has holes (14 a) regularly distributed in the first layer (14), the hole ratio being between 30% and 50%.
  4. 4. The component according to any one of the preceding claims, wherein the first layer (14) further comprises a fluoropolymer material.
  5. 5. The component according to the preceding claim, characterized in that the fluoropolymer is polytetrafluoroethylene.
  6. 6. The component according to any one of the preceding claims, wherein the fire protection barrier (13) further comprises: -a perforated third layer (16), said perforated third layer comprising glass fibres, and -A fourth layer (17) comprising fibres other than glass fibres, the third layer (16) being located between the second layer (15) and the fourth layer (17).
  7. 7. The component according to the preceding claim, characterized in that the third layer (16) has holes (16 a) regularly distributed in the third layer (16), the holes (14 a) in the first layer (14) being offset with respect to the holes (16 a) in the third layer (16).
  8. 8. The component according to any of the preceding claims, wherein the fire barrier (13) further comprises an encapsulation layer (18) attached to the body (12), the first and second layers (14, 15) being located between the encapsulation layer (18) and the body (12).
  9. 9. The component according to the preceding claim, characterized in that the envelope layer (18) has a central portion (18 a) and two connection flanges (18 b,18 c) attached to the body (12), the first and second layers (14, 15) being located between the central portion (18 a) and the body (12).
  10. 10. The component according to claim 8 or 9, characterized in that the envelope layer (18) comprises woven fibres.

Description

Component comprising a fire barrier Technical Field The present invention relates to the technical field of components comprising an organic matrix composite body and a fire barrier. The invention relates in particular to aircraft turbine components. Background Composite materials include organic matrix composite materials, also known as CMO. These composites typically comprise an organic matrix and fibers embedded in the organic matrix. Such a composite material has good mechanical strength and low density compared to a metallic material such as steel. In view of these advantageous properties, these materials are increasingly favored in the aerospace industry. Aircraft turbines are equipped with components made of organic matrix composite materials. Aircraft turbines generally extend around and along a longitudinal axis. The aircraft turbine comprises, from upstream to downstream in the direction of gas flow along the longitudinal axis, an upstream region (also referred to as fan region) in which the fan is arranged, and a downstream region in which the low-pressure compressor, the high-pressure compressor, the at least one combustion chamber, the high-pressure turbine and the low-pressure turbine are arranged. The downstream zone is a zone where the temperature may reach several hundred degrees or even up to 1500 degrees. The components (e.g., the housing) located in this downstream region must be able to withstand such high temperatures. Both the fan zone and the downstream zone are fire hazard zones where a fire may spread. This means that the components located in the fan area must also withstand high temperatures. Components comprising bodies made of organic matrix composite materials are not resistant to high temperatures and such materials may rapidly degrade in environments subjected to high temperatures (e.g., downstream regions). Furthermore, due to the low thermal resistance of such components, additional safety devices are required when such components are used in areas of the turbine where there is a fire hazard. In this case, it has been proposed that these components should be equipped with a fire barrier to protect the composite body from thermal degradation and enable implementation in downstream or fan areas. Document FR-B1-3016187 proposes a component, in particular a fan casing, having a composite body comprising a matrix and fibres embedded in the matrix. According to this document, the component also comprises a fire barrier comprising a glass fibre web pre-impregnated with a polymeric resin. For example, the resin is an epoxy resin, a phenolic resin, or a cyanate resin. The fiberglass mesh embedded in the resin forms a sacrificial refractory layer. For example, when the temperature rises during a fire, the resin may degrade, creating a thermally insulating air gap. Thus, degradation of the resin protects the body of the component from thermal attack. While this type of fire protection barrier helps limit damage to the body of the composite component when the temperature increases, this is not entirely satisfactory. Such fire barriers need to be very thick to limit thermal damage to the composite body, which tends to increase the weight and overall size of the component. It is therefore desirable to provide a component comprising a body made of an organic matrix composite material, which component can be implemented in environments subject to fire and/or high temperature risks, and which component is light in weight, compact, easy to manufacture and inexpensive. Disclosure of Invention To this end, the invention proposes a component for an aircraft turbomachine, comprising: -a body comprising a composite material comprising an organic matrix and fibers embedded in the matrix, and -A fire barrier arranged on at least one surface of the body. The component is characterized in that the fire barrier is multi-layered and comprises: a perforated first layer comprising glass fibres, and -A second layer comprising fibres other than glass fibres, the first layer being located between the body and the second layer. Due to the first and second layers of the fire barrier, and in particular due to the different properties of the fibers in the first and second layers, the first and second layers undergo different thermal expansions upon the occurrence of a fire and/or a sharp increase in temperature, resulting in delamination of the first and second layers at the interface of the first and second layers. This delamination creates an air gap between the first and second layers, providing thermal insulation to the body of the component, thus slowing degradation of the component, for example, in the event of a fire. Furthermore, due to the perforated nature of the first layer, delamination between the first layer and the second layer is located in the unperforated interface of the first layer. This limits the risk of damaging the fire barrier due to complete delamination of the first and second layers